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WO2021037047A1 - 一种飞行器的偏航角修正方法、装置及飞行器 - Google Patents

一种飞行器的偏航角修正方法、装置及飞行器 Download PDF

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Publication number
WO2021037047A1
WO2021037047A1 PCT/CN2020/111310 CN2020111310W WO2021037047A1 WO 2021037047 A1 WO2021037047 A1 WO 2021037047A1 CN 2020111310 W CN2020111310 W CN 2020111310W WO 2021037047 A1 WO2021037047 A1 WO 2021037047A1
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WO
WIPO (PCT)
Prior art keywords
yaw angle
aircraft
angle
deviation
yaw
Prior art date
Application number
PCT/CN2020/111310
Other languages
English (en)
French (fr)
Inventor
张添保
Original Assignee
深圳市道通智能航空技术有限公司
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Filing date
Publication date
Application filed by 深圳市道通智能航空技术有限公司 filed Critical 深圳市道通智能航空技术有限公司
Publication of WO2021037047A1 publication Critical patent/WO2021037047A1/zh
Priority to US17/652,007 priority Critical patent/US20220178697A1/en

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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
    • G01C25/005—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass initial alignment, calibration or starting-up of inertial devices
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G05D1/0858—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft specially adapted for vertical take-off of aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U10/00—Type of UAV
    • B64U10/10—Rotorcrafts
    • B64U10/13—Flying platforms
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U2101/00—UAVs specially adapted for particular uses or applications
    • B64U2101/30—UAVs specially adapted for particular uses or applications for imaging, photography or videography
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U2201/00—UAVs characterised by their flight controls
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U2201/00—UAVs characterised by their flight controls
    • B64U2201/10—UAVs characterised by their flight controls autonomous, i.e. by navigating independently from ground or air stations, e.g. by using inertial navigation systems [INS]
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U30/00—Means for producing lift; Empennages; Arrangements thereof
    • B64U30/20—Rotors; Rotor supports

Definitions

  • This application relates to the technical field of aircraft, and in particular to a method and device for correcting the yaw angle of an aircraft, and an aircraft.
  • Air vehicles such as Unmanned Aerial Vehicles (UAV), also known as UAVs
  • UAV Unmanned Aerial Vehicles
  • UAVs Unmanned Aerial Vehicles
  • the various actions (or attitudes) of the unmanned aerial vehicle are usually realized by controlling the different rotation speeds of multiple driving motors in the power system of the unmanned aerial vehicle.
  • the yaw angle is an important parameter in the control of the flight attitude of the unmanned aerial vehicle. That is, the yaw angle fusion of the unmanned aerial vehicle is particularly important for the attitude control of the unmanned aerial vehicle.
  • the unmanned aerial vehicle cannot fly in the preset direction or trajectory. In the worst case, the phenomenon of panning may occur, and the aircraft may even become unstable and blow up.
  • the magnetic interference of the aircraft in the indoor environment is serious, and the GPS information is poor.
  • the aircraft is flying or hovering indoors. Since there is no GPS information correction, the magnetometer is also seriously disturbed. There is no available information to correct the yaw angle.
  • the instrument integral itself has drift characteristics, so when flying or hovering indoors, the yaw angle of the aircraft is prone to drift.
  • the embodiments of the present invention provide an aircraft yaw angle correction method, device, and aircraft, which solve the problem that indoor aircraft rely on visual information for yaw angle correction and indoor magnetic interference affects yaw angle correction, and improve the aircraft flying or hanging indoors. Stop the stability.
  • the embodiments of the present invention provide the following technical solutions:
  • an embodiment of the present invention provides an aircraft yaw angle correction method, which is applied to the aircraft, and the method includes:
  • IMU data includes IMU acceleration information and IMU angular velocity information
  • a fusion yaw angle is generated.
  • the determining the yaw angle of the magnetometer according to the magnetometer data includes:
  • the magnetometer data of the standard magnetic field of the aircraft at the current position is compared to calculate the magnetometer yaw angle.
  • the determining the initial value of the yaw angle according to the magnetometer yaw angle includes:
  • the determining the yaw angular velocity compensation amount according to the magnetometer data includes:
  • the determining the yaw angle deviation angle according to the magnetometer yaw angle includes:
  • the determining the relative deviation angle of the yaw angle according to the deviation angle of the yaw angle includes:
  • the relative deviation angle of the yaw angle is determined according to the deviation angle of the yaw angle, the ground height of the aircraft, and the flying height of the aircraft.
  • the determining the relative deviation angle of the yaw angle according to the yaw angle deviation angle, the ground height of the aircraft, and the flying height of the aircraft includes:
  • the determining the relative offset of the yaw angle deviation according to the ground height of the aircraft and the flying height of the aircraft includes:
  • the altitude of the aircraft meets the first preset condition
  • the derivative of the deviation angle of the yaw angle satisfies the third preset condition.
  • the first preset condition is:
  • the height of the aircraft to the ground is greater than 0.4m, and the duration is not less than 0.5s.
  • the second preset condition is:
  • the flying height of the aircraft is greater than 0.4m, and the duration is not less than 0.5s.
  • the third preset condition is:
  • the absolute value of the differential of the yaw angle deviation angle is less than 0.1 and the duration is not less than 0.5s.
  • the determining the relative deviation angle of the yaw angle according to the deviation angle of the yaw angle and the relative deviation amount of the yaw angle deviation includes:
  • the relative compensation value of the yaw angle error is determined to be the relative deviation angle of the yaw angle:
  • the relative compensation value of the yaw angle error satisfies the fourth preset condition, wherein the relative compensation value of the yaw angle error is the difference between the yaw angle deviation angle and the relative offset of the yaw angle deviation.
  • the fourth preset condition is:
  • the absolute value of the relative compensation value of the yaw angle error is less than 0.1 and the duration is not less than 0.5s.
  • the determining the yaw angular velocity compensation amount according to the relative deviation angle of the yaw angle includes:
  • a feedback control algorithm is used to calculate the relative deviation angle of the yaw angle to determine the yaw angular velocity compensation amount.
  • an embodiment of the present invention provides an aircraft yaw angle correction device, which is applied to the aircraft, and the device includes:
  • the fusion yaw angle generation module is used to generate a fusion yaw angle according to the initial value of the yaw angle and the relative value of the yaw angle.
  • the determination module includes a calibration and coordinate system conversion module, and the calibration and coordinate system conversion module is used to:
  • the magnetometer data of the standard magnetic field of the aircraft at the current position is compared to calculate the magnetometer yaw angle.
  • the determination module further includes a static state detection module, and the static state detection module is configured to:
  • the determination module further includes a yaw angle deviation judgment and processing module, and the yaw angle deviation judgment and processing module is used for:
  • the yaw angle deviation determination and processing module is used to:
  • the yaw angle deviation determination and processing module is used to:
  • the relative deviation angle of the yaw angle is determined according to the deviation angle of the yaw angle, the ground height of the aircraft, and the flying height of the aircraft.
  • the yaw angle deviation judgment and processing module includes a logical OR operation module, and the logical OR operation module is used for:
  • the altitude of the aircraft meets the first preset condition
  • the derivative of the deviation angle of the yaw angle satisfies the third preset condition.
  • the first preset condition is:
  • the height of the aircraft to the ground is greater than 0.4m, and the duration is not less than 0.5s.
  • the second preset condition is:
  • the flying height of the aircraft is greater than 0.4m, and the duration is not less than 0.5s.
  • the third preset condition is:
  • the absolute value of the differential of the yaw angle deviation angle is less than 0.1 and the duration is not less than 0.5s.
  • the yaw angle deviation determination and processing module includes a logical AND operation module, and the logical AND operation module is used for:
  • the relative compensation value of the yaw angle error is determined to be the relative deviation angle of the yaw angle:
  • the relative compensation value of the yaw angle error satisfies the fourth preset condition, wherein the relative compensation value of the yaw angle error is the difference between the yaw angle deviation angle and the relative offset of the yaw angle deviation.
  • the fourth preset condition is:
  • the absolute value of the relative compensation value of the yaw angle error is less than 0.1 and the duration is not less than 0.5s.
  • the determining module includes a feedback control module, and the feedback control module is configured to:
  • a feedback control algorithm is used to calculate the relative deviation angle of the yaw angle to determine the yaw angular velocity compensation amount.
  • an embodiment of the present invention provides an aircraft, including:
  • An arm connected to the fuselage
  • a power device which is provided on the fuselage and/or the arm, and is used to provide power for the aircraft to fly;
  • the flight controller is located on the fuselage
  • the flight controller includes:
  • At least one processor and,
  • a memory communicatively connected with the at least one processor; wherein,
  • the memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor, so that the at least one processor can execute the method for correcting the yaw angle of the aircraft as described above .
  • an embodiment of the present invention also provides a non-volatile computer-readable storage medium, the computer-readable storage medium stores computer-executable instructions, and the computer-executable instructions are used to enable the aircraft to execute the above The method for correcting the yaw angle of the aircraft.
  • the invention can solve the problem that the indoor aircraft relies on visual information to perform yaw angle correction and indoor magnetic interference affects the yaw angle correction, and improves the stability of the aircraft flying or hovering indoors.
  • Figure 1 is a specific structural diagram of an aircraft provided by an embodiment of the present invention.
  • FIG. 2 is a schematic block diagram of a method for correcting yaw angle of an aircraft according to an embodiment of the present invention
  • FIG. 3 is a schematic diagram of a yaw angle deviation judgment and processing algorithm provided by an embodiment of the present invention.
  • FIG. 4 is a schematic flowchart of a method for correcting yaw angle of an aircraft according to an embodiment of the present invention
  • FIG. 5 is a detailed flowchart of step S20 in FIG. 4;
  • FIG. 6 is a detailed flowchart of step S30 in FIG. 4;
  • FIG. 7 is a detailed flowchart of step S40 in FIG. 4;
  • FIG. 8 is a detailed flowchart of step S42 in FIG. 7;
  • FIG. 9 is a detailed flowchart of step S421 in FIG. 8;
  • FIG. 10 is a detailed flowchart of step S423 in FIG. 8;
  • FIG. 11 is a schematic structural diagram of an aircraft yaw angle correction device provided by an embodiment of the present invention.
  • FIG. 12 is a schematic diagram of the structure of the determining module in FIG. 11;
  • FIG. 13 is a schematic structural diagram of the yaw angle deviation judgment and processing module in FIG. 12;
  • FIG. 14 is a schematic diagram of the hardware structure of an aircraft provided by an embodiment of the present invention.
  • 15 is a connection block diagram of an aircraft provided by an embodiment of the present invention.
  • Fig. 16 is a schematic diagram of the power system in Fig. 15.
  • the method for correcting the yaw angle of the aircraft can be applied to various movable objects driven by motors or motors, including but not limited to aircraft, robots, and the like.
  • the aircraft may include unmanned aerial vehicles (UAV), unmanned aerial vehicles, and so on.
  • UAV unmanned aerial vehicles
  • the method for correcting the yaw angle of the aircraft in the embodiment of the present invention is applied to the flight controller of the aircraft.
  • FIG. 1 is a specific structure diagram of an aircraft provided by an embodiment of the present invention.
  • the aircraft 10 includes: a fuselage 11, an arm 12 connected to the fuselage 11, a power device 13 provided on the arm 12, and a pan/tilt 14 connected to the bottom of the fuselage 11 , A camera 15 installed on the pan-tilt 14 and a flight controller (not shown) installed in the fuselage 11.
  • the flight controller is connected to the power device 13, and the power device 13 is installed on the fuselage 11 to provide flight power for the aircraft 10.
  • the flight controller is used to execute the above-mentioned method for correcting the yaw angle of the aircraft to correct the yaw angle of the aircraft, and generate a control command according to the yaw angle of the fused aircraft, and send the control command to the power unit 13
  • the ESC controls the driving motor of the power unit 13 through the control command.
  • the flight controller is used to execute the yaw angle correction method of the aircraft to correct the yaw angle of the aircraft, and send the corrected yaw angle of the aircraft to the ESC, and the ESC generates control according to the corrected yaw angle of the aircraft Command and control the drive motor of the power unit 13 through the control command.
  • the fuselage 11 includes a central shell and one or more arms connected to the central shell, and the one or more arms extend radially from the central shell.
  • the connection between the arm and the center housing can be an integral connection or a fixed connection.
  • the power unit is installed on the arm.
  • the flight controller is used to execute the above-mentioned aircraft yaw angle correction method to correct the yaw angle of the aircraft, and generate a control command according to the corrected yaw angle of the aircraft, and send the control command to the ESC of the power unit for electrical
  • the drive motor of the power plant is controlled by the control command.
  • the controller is a device with a certain logic processing capability, such as a control chip, a single-chip microcomputer, and a microcontroller unit (MCU).
  • the power unit 13 includes: an ESC, a drive motor and a propeller.
  • the ESC is located in the cavity formed by the arm or the center housing.
  • the ESC is connected to the controller and the drive motor respectively.
  • the ESC is electrically connected to the drive motor, and is used to control the drive motor.
  • the driving motor is installed on the arm, and the rotating shaft of the driving motor is connected to the propeller.
  • the propeller generates a force for moving the aircraft 10 under the driving of the driving motor, for example, a lift force or a thrust force for moving the aircraft 10.
  • the completion of various prescribed speeds and actions (or attitudes) of the aircraft 10 is achieved by controlling the drive motor through an ESC.
  • the full name of the ESC is electronic governor, which adjusts the rotation speed of the driving motor of the aircraft 10 according to the control signal.
  • the controller is the executive body that executes the method for correcting the yaw angle of the aircraft, and the ESC controls the driving motor based on the control command generated by the yaw angle of the fused aircraft.
  • the principle of ESC to control the drive motor is roughly as follows: the drive motor is an open-loop control element that converts electrical pulse signals into angular displacement or linear displacement.
  • the speed and stop position of the drive motor depends only on the frequency and pulse number of the pulse signal, and is not affected by the load change.
  • the drive receives a pulse signal, it drives the drive motor of the power unit Rotate a fixed angle according to the set direction, and its rotation runs at a fixed angle. Therefore, the ESC can control the angular displacement by controlling the number of pulses, so as to achieve the purpose of accurate positioning; at the same time, the speed and acceleration of the driving motor can be controlled by controlling the pulse frequency, so as to achieve the purpose of speed regulation.
  • the main functions of the aircraft 10 are aerial photography, real-time image transmission, and detection of high-risk areas.
  • the aircraft 10 will be connected with a camera component.
  • the aircraft 10 and the camera assembly are connected by a connecting structure, such as a vibration damping ball.
  • the camera component is used to obtain a photographed image during the aerial photographing process of the aircraft 10.
  • the camera component includes: a pan-tilt and a camera.
  • the gimbal is connected to the aircraft 10.
  • the photographing device is mounted on the pan-tilt.
  • the photographing device may be an image acquisition device for collecting images.
  • the photographing device includes but is not limited to a camera, a video camera, a camera, a scanner, a camera phone, and the like.
  • the pan/tilt is used to mount the camera to realize the fixation of the camera, or adjust the posture of the camera at will (for example, change the height, inclination and/or direction of the camera) and keep the camera stably in the set posture on.
  • the pan-tilt is mainly used to stably maintain the camera in a set posture, to prevent the camera from shaking and ensure the stability of the camera.
  • the pan-tilt 14 is connected with the flight controller to realize data interaction between the pan-tilt 14 and the flight controller.
  • the flight controller sends a yaw command to the gimbal 14.
  • the gimbal 14 obtains and executes the speed and direction command of the yaw, and sends the data information generated after the yaw command is executed to the flight controller for the flight controller Detect the current yaw status.
  • PTZ includes: PTZ motor and PTZ base.
  • the gimbal motor is installed on the base of the gimbal.
  • the flight controller can also control the gimbal motor through the ESC of the power unit 13. Specifically, the flight controller is connected to the ESC, and the ESC is electrically connected to the gimbal motor.
  • the flight controller generates the gimbal motor control command, and the ESC passes PTZ motor control commands to control the PTZ motor.
  • the gimbal base is connected with the fuselage of the aircraft, and is used to fix the camera assembly on the fuselage of the aircraft.
  • the pan/tilt motor is respectively connected with the pan/tilt base and the camera.
  • the pan/tilt can be a multi-axis pan/tilt. To adapt to it, there are multiple pan/tilt motors, that is, one pan/tilt motor is provided for each axis. On the one hand, the pan/tilt motor can drive the rotation of the shooting device, so as to meet the adjustment of the horizontal rotation and pitch angle of the shooting shaft.
  • the rotation of the pan/tilt motor By manually remotely controlling the rotation of the pan/tilt motor or using a program to make the motor rotate automatically, it can achieve the function of omnidirectional scanning and monitoring;
  • the rotation of the pan/tilt motor cancels the disturbance of the camera in real time, prevents the camera from shaking, and ensures the stability of the shooting image.
  • the camera is mounted on the pan/tilt.
  • the camera is equipped with an inertial measurement unit (IMU).
  • IMU inertial measurement unit
  • the inertial measurement unit is a device for measuring the three-axis attitude angle (or angular velocity) and acceleration of an object.
  • a three-axis gyroscope and three-directional accelerometers are installed in an IMU to measure the angular velocity and acceleration of the object in three-dimensional space, and to calculate the posture of the object.
  • the IMU should be installed on the center of gravity of the aircraft.
  • the yaw angle of the aircraft is an important parameter in controlling the attitude of the aircraft, and the drive motor needs to be controlled based on the yaw angle of the aircraft.
  • the yaw angle of the aircraft is obtained in real time through the aircraft controller, and the necessary attitude information is provided for the attitude control of the aircraft. That is to say, the correct estimation of the aircraft's yaw angle is particularly important for the attitude control of the aircraft. If the aircraft's yaw angle is estimated incorrectly, the aircraft may not be able to fly in the preset direction or trajectory at a low level, and may become unstable and cause the aircraft to blow up.
  • the magnetometer In the indoor environment, because there is no GPS information correction, the magnetometer is also seriously interfered, which leads to the lack of sufficient available information to correct the yaw angle. Moreover, due to the drift characteristics of the gyroscope integral itself, it is indoors. When flying or hovering, the aircraft is prone to yaw angle deviation.
  • the main purpose of the embodiments of the present invention is to provide a method, device and aircraft for correcting the yaw angle of an aircraft, which can correct the yaw angle of the aircraft based on IMU data and magnetometer data to solve the problem of indoor aircraft Relying on visual information to correct the yaw angle and the problem of indoor magnetic interference affecting the yaw angle correction, thereby improving the stability of the aircraft flying or hovering indoors.
  • the embodiment of the present invention obtains IMU data and magnetometer data, calculates the initial value of the yaw angle and the relative value of the yaw angle, and fuses the initial value of the yaw angle and the relative value of the yaw angle.
  • the fusion method can avoid indoor magnetic Interference, that is, in the absence of GPS signals and strong magnetic interference, the stability of the flight or hover of the aircraft can also be ensured.
  • FIG. 2 is a schematic block diagram of a method for correcting yaw angle of an aircraft according to an embodiment of the present invention
  • the aircraft is stationary, and the stationary flag signal is generated.
  • the stationary flag signal is input to the enabling module of the aircraft, and the magnetometer yaw angle output is used as the initial value of the yaw angle according to the rising edge of the stationary flag signal, and the magnetometer data is transformed by standard matrix rotation to generate Magnetometer yaw angle, according to the magnetometer yaw angle and the current fusion yaw angle, calculate the yaw angle deviation angle, judge and process the yaw angle deviation angle, and generate the relative deviation angle of the yaw angle.
  • the yaw angular velocity compensation amount is generated, and the corrected acceleration is generated by fusing the IMU angular velocity and the yaw angular velocity compensation amount generated according to the magnetometer data.
  • Perform integration to obtain the relative value of the yaw angle, and fuse the initial value of the yaw angle and the relative value of the yaw angle to generate a fused yaw angle.
  • FIG. 3 is a schematic diagram of a yaw angle deviation judgment and processing algorithm provided by an embodiment of the present invention.
  • FIG. 4 is a schematic flowchart of a method for correcting yaw angle of an aircraft according to an embodiment of the present invention
  • the yaw angle correction method of the aircraft can be executed by various electronic devices with certain logic processing capabilities, such as aircraft, control chips, etc.
  • the aircraft can include unmanned aerial vehicles, unmanned ships, and the like.
  • the following electronic equipment takes an aircraft as an example for description.
  • the aircraft is connected with a gimbal, which includes a gimbal motor and a gimbal base.
  • the gimbal can be a multi-axis gimbal, such as a two-axis gimbal, a three-axis gimbal, and the following three-axis gimbal as an example. Description.
  • the specific structure of the aircraft and the gimbal reference may be made to the above description, and therefore, it will not be repeated here.
  • the method is applied to an aircraft, such as an unmanned aerial vehicle, and the method includes:
  • Step S10 Obtain IMU data and magnetometer data, where the IMU data includes IMU acceleration information and IMU angular velocity information;
  • the aircraft is provided with an attitude sensor assembly
  • the attitude sensor assembly includes: an inertial measurement unit (IMU), a magnetometer, etc.
  • IMU inertial measurement unit
  • the magnetometer is used
  • the inertial measurement unit includes a gyroscope and an accelerometer
  • the gyroscope is used to obtain IMU angular velocity
  • the accelerometer is used to obtain IMU acceleration information
  • the IMU data includes: IMU acceleration information and IMU Angular velocity information
  • the magnetometer data includes: magnetic field intensity information.
  • the IMU data is acquired through the inertial measurement unit, and the IMU data is calibrated and coordinate system conversion is performed to generate IMU acceleration information and IMU angular velocity information, where the IMU acceleration information is the measurement data of the inertial measurement unit after the calibration matrix After calibration and the coordinate transformation from the body coordinate system to the ground coordinate system, the acceleration information in the ground coordinate system is obtained.
  • the calibration matrix is calibrated by the user at the place where the user wants to fly.
  • the calibration matrix is different anywhere on the earth.
  • the aircraft reports magnetometer interference, and the user is required to calibrate the calibration matrix before determining the calibration matrix.
  • the conversion of the airframe coordinate system to the ground coordinate system is completed by a rotation transformation matrix.
  • a rotation transformation matrix is generated according to the attitude angle of the aircraft, and the IMU data is converted from the airframe coordinates through the rotation transformation matrix.
  • the system is converted to the ground coordinate system, and the IMU acceleration information and the IMU angular velocity information are generated.
  • the attitude angle of the aircraft includes: a yaw angle, a pitch angle, and a roll angle, where the yaw angle is the current fused yaw angle, that is, the real-time fused yaw angle will be used to calculate the rotation transformation matrix , And then used for the next fusion to continuously update the fusion yaw angle.
  • the rotation transformation matrix is a 3*3 matrix, which contains the sine and cosine functions of the yaw angle, pitch angle, and roll angle, and different functions are selected according to the specific situation.
  • the rotation transformation matrix is:
  • ( ⁇ , ⁇ , ⁇ ) is the attitude angle
  • ⁇ is the roll angle in the attitude angle
  • ⁇ is the pitch angle in the attitude angle
  • ⁇ is the yaw angle in the attitude angle
  • Step S20 Determine the yaw angle of the magnetometer according to the magnetometer data
  • the magnetometer data is obtained by a magnetometer
  • the magnetometer data includes: magnetic field intensity information
  • the magnetic field intensity is a triaxial magnetic field intensity
  • the magnetometer data measured by the magnetometer is a triaxial magnetic field in the body coordinate system Therefore, it is necessary to remove bias and cross-coupling through the calibration matrix, and transform it to the ground coordinate system through the rotation matrix.
  • FIG. 5 is a detailed flowchart of step S20 in FIG. 4;
  • the determining the yaw angle of the magnetometer according to the magnetometer data includes:
  • Step S21 Calibrate the magnetometer data to generate calibrated magnetometer data
  • the magnetometer data is calibrated according to a preset calibration matrix to generate calibrated magnetometer data; specifically, the preset calibration matrix is calibrated by the user at the place where the user wants to fly, and the calibration matrix is on the earth Any place above is different.
  • the aircraft reports magnetometer interference and requires the user to calibrate before determining the calibration matrix.
  • Step S22 Obtain the attitude angle of the aircraft and generate a rotation transformation matrix according to the attitude angle of the aircraft;
  • the rotation transformation matrix is used to convert the airframe coordinate system into a ground coordinate system
  • the attitude angle of the aircraft includes: a yaw angle, a pitch angle, and a roll angle.
  • the yaw angle is the current fusion yaw angle, that is, the real-time fusion yaw angle will be used to calculate the rotation transformation matrix, and then used for the next fusion, to continuously update the fusion yaw angle.
  • the rotation transformation matrix is a 3*3 matrix, which contains the sine and cosine functions of the yaw angle, pitch angle, and roll angle, and different functions are selected according to the specific situation. Generally speaking, by first rotating the yaw angle , Then turn the pitch angle, and finally turn the roll angle.
  • the rotation transformation matrix is:
  • ( ⁇ , ⁇ , ⁇ ) is the attitude angle
  • ⁇ is the roll angle in the attitude angle
  • ⁇ is the pitch angle in the attitude angle
  • ⁇ is the yaw angle in the attitude angle
  • Step S23 Use the rotation transformation matrix to perform coordinate transformation on the calibrated magnetometer data to generate magnetometer data in the ground coordinate system
  • the calibrated magnetometer data that is, the magnetic field strength
  • the rotation transformation matrix to generate the magnetic field strength in the ground coordinate system, which is equivalent to performing the body coordinate system on the magnetometer data
  • the magnetometer data in the ground coordinate system is generated after coordinate transformation.
  • Step S24 According to the magnetometer data in the ground coordinate system, compare the magnetometer data of the standard magnetic field of the aircraft at the current position to calculate the magnetometer yaw angle.
  • the current position of the aircraft corresponds to a standard magnetic field
  • the three-axis readings of the magnetometer form a vector
  • the magnetometer data of the standard magnetic field at the current position corresponds to a vector.
  • the gauge data is compared with the magnetometer data of the standard magnetic field at the current position of the aircraft, the vector angle between the two is calculated, and the vector angle is used as the magnetometer yaw angle.
  • Step S30 Determine the initial value of the yaw angle according to the yaw angle of the magnetometer
  • the aircraft is provided with an enabling module, the enabling module includes an input terminal and an output terminal, when the enabling module receives the rising edge of the stationary flag signal, the enabling module transfers the magnetic force
  • the yaw angle is calculated and output as the initial value of the yaw angle.
  • FIG. 6 is a detailed flowchart of step S30 in FIG. 4;
  • the determining the initial value of the yaw angle according to the yaw angle of the magnetometer includes:
  • Step S31 Determine whether the aircraft has changed from a stationary state to a moving state at the current moment
  • the signal of the stationary flag of the aircraft is determined according to the stationary state of the aircraft; wherein the stationary flag of the aircraft is used to characterize the stationary state of the aircraft, and the stationary state of the aircraft is determined according to the stationary state of the aircraft.
  • the signal for determining the stationary flag of the aircraft includes: acquiring IMU acceleration and IMU angular velocity in the IMU data, and determining the stationary state of the aircraft by performing stationary detection on the IMU acceleration and IMU angular velocity. If the stationary state of the aircraft is stationary, the value of the stationary state bit is 1. If the stationary state of the aircraft is moving, the value of the stationary state bit is 0. It can be understood that the IMU is at rest.
  • the data of acceleration and IMU angular velocity fluctuate particularly small and stable.
  • the method specifically includes: transforming the body coordinate system and the ground coordinate system on the IMU data according to the rotation transformation matrix to generate the IMU acceleration in the ground coordinate system and the IMU angular velocity in the ground coordinate system;
  • the IMU acceleration in the coordinate system and the IMU angular velocity in the ground coordinate system determine the stationary state of the aircraft, and generate the stationary flag of the aircraft.
  • Step S32 If yes, use the magnetometer yaw angle as the initial value of the yaw angle.
  • the signal of the stationary state bit is input to the enabling module of the aircraft, and when the enabling module detects that the signal of the stationary state bit has a rising edge, the magnetometer yaw angle output is used as the The initial value of the yaw angle.
  • the signal of the stationary state bit is 0 when the aircraft is stationary and 1 when the aircraft is in motion.
  • the signal of the stationary state bit is input to the enabling module of the aircraft, and when the enabling module detects the The signal of the stationary state bit has a rising edge, that is, when the stationary state bit is from 0 to 1, the enabling module uses the magnetometer yaw angle output as the initial value of the yaw angle.
  • Step S40 Determine the yaw angular velocity compensation amount according to the magnetometer data
  • the IMU data measured by the IMU needs to be corrected.
  • the yaw angular velocity compensation amount is used to correct the IMU angular velocity obtained by the IMU, and the yaw angle of the magnetometer needs to be used to determine the yaw angle.
  • Angular velocity compensation amount for details, please refer to Fig. 7, which is a detailed flowchart of step S40 in Fig. 4;
  • the determining the yaw angular velocity compensation amount according to the magnetometer yaw angle includes:
  • Step S41 Determine the deviation angle of the yaw angle according to the yaw angle of the magnetometer
  • the determining the yaw angle deviation angle according to the magnetometer yaw angle includes: determining the yaw angle deviation angle according to the magnetometer yaw angle and the fused yaw angle at the previous time.
  • the current fusion yaw angle is a real-time fusion yaw angle
  • the difference between the current fusion yaw angle and the magnetometer yaw angle is used as the yaw angle deviation angle.
  • Step S42 Determine the relative deviation angle of the yaw angle according to the deviation angle of the yaw angle
  • the determining the relative deviation angle of the yaw angle according to the deviation angle of the yaw angle includes:
  • the relative deviation angle of the yaw angle is determined according to the deviation angle of the yaw angle, the ground height of the aircraft, and the flying height of the aircraft.
  • the determining the relative deviation angle of the yaw angle according to the deviation angle of the yaw angle, the ground height of the aircraft, and the flying height of the aircraft includes:
  • the determining the relative offset of the yaw angle deviation according to the ground height of the aircraft and the flying height of the aircraft includes:
  • the altitude of the aircraft meets the first preset condition
  • the derivative of the deviation angle of the yaw angle satisfies the third preset condition.
  • the first preset condition is: the altitude of the aircraft above the ground is greater than 0.4m, and the duration is not less than 0.5s
  • the second preset condition is: the altitude of the aircraft is greater than 0.4m, and The duration is not less than 0.5s
  • the third preset condition is: the absolute value of the derivative of the yaw angle deviation angle is less than 0.1 and the duration is not less than 0.5s.
  • the relative deviation angle of the yaw angle is the relative deviation amount of the yaw angle after the yaw angle deviation is determined and processed.
  • FIG. 8 is a detailed flowchart of step S42 in FIG. 7;
  • the determining the relative deviation angle of the yaw angle according to the deviation angle of the yaw angle includes:
  • Step S421 Determine the relative offset of the yaw angle deviation according to the yaw angle deviation angle
  • the aircraft is provided with a latch module, and the latch module is used to latch the yaw angle deviation angle, and when a valid signal is received, the latch module outputs the yaw angle deviation angle As the yaw angle deviation relative offset, the yaw angle deviation angle is continuously latched, and the yaw angle deviation relative offset is not updated until the next valid signal arrives.
  • FIG. 9 is a detailed flowchart of step S421 in FIG. 8;
  • the determining the relative offset of the yaw angle deviation according to the yaw angle deviation angle includes:
  • Step S4211 Obtain the height reset pulse signal and the differential reset pulse signal
  • the obtaining the height reset pulse includes:
  • the altitude of the aircraft refers to the distance from the position of the aircraft when the aircraft is flying to the ground directly below
  • the flying height of the aircraft refers to the height difference between the position of the aircraft during the flight and the take-off point.
  • the aircraft is provided with an ultrasonic sensor and/or a TOF sensor (Time of Flight, TOF), and the ground height and flying height of the aircraft are acquired by the ultrasonic sensor and/or TOF sensor.
  • TOF Time of Flight
  • the height determination flag signal is used to avoid the magnetic interference of the reinforced concrete in the floor.
  • the ground height threshold and its duration threshold are preset, if the ground height of the aircraft is greater than the ground height threshold , And the duration of the ground height greater than the ground height threshold is greater than the duration threshold, the height determination flag signal outputs a high level, otherwise, the height determination flag signal outputs a low level Or, preset the flight altitude threshold and its duration threshold, if the flight altitude of the aircraft is greater than the flight altitude threshold, and the flight altitude of the aircraft is greater than the flight altitude threshold for a duration greater than the duration threshold , The height determination flag signal outputs a high level, otherwise, the height determination flag outputs a low level, for example: the preset ground height threshold and its duration threshold are 0.4m and 0.5s, respectively, The preset flight altitude threshold and its duration threshold are 0.4m and 0.5s, respectively.
  • the altitude determination flag signal When the aircraft's ground altitude is greater than 0.4m and lasts for more than 0.5s, the altitude determination flag signal outputs 1; or, when the aircraft's altitude is greater than 0.4m When it lasts for more than 0.5s, the height judgment flag signal outputs 1; otherwise, the height judgment flag signal outputs 0.
  • the ground height threshold and its duration threshold in the embodiments of the present invention, as well as the flying height threshold and its duration threshold can be specifically set according to specific requirements, and all fall within the protection scope of the present invention.
  • an altitude reset pulse signal is generated.
  • the controller of the aircraft is provided with a rising edge detection module.
  • the rising edge detection module acquires the height determination flag signal and determines whether the height determination flag signal has a rising edge, and if so, generates the height Reset pulse signal.
  • the obtaining the differential reset pulse signal includes:
  • the yaw angle deviation angle is differentiated and filtered, the initial yaw angle differential value is obtained by differentiating the yaw angle, and the initial yaw angle differential value is filtered,
  • the final yaw angle differential value that is, the yaw angle differential value is generated, wherein the filtering process includes low-pass filtering and Kalman filtering, which are respectively completed by the low-pass filter and Kalman filter of the aircraft.
  • the controller of the aircraft is provided with a judgment logic module, the judgment logic module obtains the yaw angle differential value, performs a logical judgment on the yaw angle differential value, and generates the yaw angle differential judgment flag Bit signal, wherein the judgment logic of the judgment logic module on the differential value of the yaw angle includes: a preset absolute value threshold of the differential value and its duration threshold, if the differential value of the yaw angle is less than the differential value Absolute value threshold, and if the duration of the yaw angle differential value less than the absolute value threshold of the differential value is greater than the duration threshold, the yaw angle differential determination flag signal outputs a high level; otherwise, The yaw angle differential judgment flag signal outputs a low level, for example: the preset absolute value threshold of the differential value is 0.1, the duration threshold value is 0.5s, if the absolute value of the yaw angle differential value is less than 0.1 And, if the duration of the absolute value of the yaw angle
  • the rising edge detection module of the aircraft obtains the yaw angle differential determination flag signal, and determines whether the yaw angle differential determination flag signal has a rising edge, and if so, generates a differential reset pulse signal , And the differential reset pulse signal outputs a high level, otherwise, the differential reset pulse signal outputs a low level.
  • Step S4212 Perform logical judgment on the height reset pulse signal and the differential reset pulse signal, and generate a yaw angle relative offset reset pulse flag signal;
  • a logical OR operation is performed on the altitude reset pulse signal and the differential reset pulse signal, and the result after the logical OR operation is used as the yaw angle relative offset reset pulse flag signal. For example, if either of the altitude reset pulse signal or the differential reset pulse signal is at a high level, the yaw angle relative offset reset pulse flag signal is at a high level, and if the altitude is reset The set pulse signal and the differential reset pulse signal are both low, and the yaw angle relative offset reset pulse flag signal is low.
  • Step S4213 Send the yaw angle relative offset reset pulse flag signal to the latch module of the aircraft, if the yaw angle relative offset reset pulse is received by the latch module of the aircraft The effective signal of the flag bit, the yaw angle deviation angle is output as the yaw angle deviation relative offset.
  • the effective signal refers to a high-level signal, namely 1, if the yaw angle relative offset reset pulse flag signal is high, the latch module of the aircraft will The yaw angle deviation angle is output as the yaw angle deviation relative offset. If the yaw angle relative offset reset pulse flag signal is low, the latch module of the aircraft will continue to latch the yaw angle deviation. The yaw angle deviation angle until receiving the valid signal of the yaw angle relative offset reset pulse flag signal, that is, the yaw angle relative offset reset pulse flag signal is a high level signal.
  • Step S422 Determine a relative compensation value of the yaw angle error according to the yaw angle deviation angle and the yaw angle deviation relative offset;
  • yaw angle error relative compensation value yaw angle deviation angle-yaw angle deviation relative offset.
  • Step S423 Determine the relative deviation angle of the yaw angle according to the relative compensation value of the yaw angle error.
  • FIG. 10 is a detailed flowchart of step S423 in FIG. 8;
  • the determining the relative deviation angle of the yaw angle according to the relative compensation value of the yaw angle error includes:
  • Step S4231 Perform a logical judgment on the relative compensation value of the yaw angle error, and generate a judgment flag signal for the relative compensation value of the yaw angle error;
  • the judgment logic module of the aircraft obtains the relative compensation value of the yaw angle error, and performs a logical judgment on the relative compensation value of the yaw angle error, wherein the judgment logic module judges the yaw angle error
  • the judgment logic of the relative compensation value includes: a preset absolute value threshold of the compensation value and its duration threshold, if the relative compensation value of the yaw angle error is less than the absolute value of the compensation threshold, and the yaw angle error is relatively
  • the compensation value is less than the absolute value threshold of the compensation value and the duration is greater than the duration threshold, the yaw angle error relative compensation value determination flag signal outputs a high level, otherwise, the yaw angle error relative compensation value It is determined that the flag signal outputs a low level.
  • the absolute value threshold of the compensation value differential value is preset to be 0.1, and the duration threshold value is 0.5s. If the absolute value of the yaw angle error relative to the compensation value is less than 0.1, And, if the duration of the relative compensation value of the yaw angle error is less than 0.1 is longer than the 0.5s, the yaw angle error relative compensation value determination flag signal outputs 1, otherwise, the yaw angle error The relative compensation value judgment flag signal outputs 0. It can be understood that the absolute value threshold of the compensation value and its duration threshold can be specifically set according to specific requirements, and both fall within the protection scope of the present invention.
  • Step S4232 Perform logical judgment on the altitude judgment flag signal, the yaw angle differential judgment flag signal, and the yaw angle error relative compensation value judgment flag signal to generate a yaw angle deviation compensation flag;
  • a logical AND operation is performed on the altitude judgment flag signal, the yaw angle differential judgment flag signal, and the yaw angle error relative compensation value judgment flag signal to determine the value of the yaw angle deviation compensation flag.
  • the yaw angle deviation compensation flag is used to determine whether the relative compensation value of the yaw angle error can be used for compensation. If the yaw angle deviation compensation flag is high, it means it can be used for compensation. , If the yaw angle deviation compensation flag is low, it means that it cannot be used for compensation.
  • the aircraft's determination logic module determines the altitude If the value of the judgment flag signal, the yaw angle differential judgment flag signal, and the yaw angle error relative compensation value judgment flag signal is high, the yaw angle deviation compensation flag output is high Level, or if any one of the altitude determination flag signal, the yaw angle differential determination flag signal, and the yaw angle error relative compensation value determination flag signal is low, the aircraft's determination logic The module performs a logical AND operation on the altitude judgment flag signal, the yaw angle differential judgment flag signal, and the yaw angle error relative compensation value judgment flag signal. The value is low, then the yaw angle deviation The compensation flag bit outputs low level.
  • Step S4233 Input the signal of the yaw angle deviation compensation flag to the enabling module of the aircraft, and if the enabling module of the aircraft receives the valid signal of the yaw angle deviation compensation flag, the The enabling module outputs the relative compensation value of the yaw angle error as the relative deviation angle of the yaw angle.
  • the enabling module of the aircraft will output the relative compensation value of the yaw angle error as the relative deviation angle of the yaw angle.
  • Step S43 Determine the yaw angular velocity compensation amount according to the relative deviation angle of the yaw angle.
  • the aircraft is provided with a feedback controller that inputs the relative deviation angle of the yaw angle into the feedback controller, and the feedback controller calculates the relative deviation angle of the yaw angle through a feedback control algorithm to determine
  • Step S50 Determine the corrected angular velocity according to the IMU angular velocity information and the yaw angular velocity compensation amount;
  • Step S60 Determine the relative value of the yaw angle according to the corrected angular velocity
  • the corrected angular velocity is integrated, and the integral value of the corrected angular velocity is used as the relative value of the yaw angle.
  • Step S70 Generate a fusion yaw angle according to the initial value of the yaw angle and the relative value of the yaw angle.
  • the error calculation is performed once for each sampling step of the aircraft, and the error calculation is performed indefinitely without stopping through the feedback loop.
  • the fused yaw angle of the fusion is to be deviated from the yaw angle of the magnetometer to generate the yaw angle error angle, which will continue endlessly until the aircraft is powered off.
  • the fusion yaw angle is also endlessly updated, and each sampling moment corresponds to a unique fusion yaw angle, and the yaw angle error angle is the difference between the magnetometer yaw angle and the fusion yaw angle .
  • an aircraft yaw angle correction method includes: acquiring IMU data and magnetometer data, wherein the IMU data includes IMU acceleration information and IMU angular velocity information; Determine the yaw angle of the magnetometer according to the magnetometer data; determine the initial value of the yaw angle according to the magnetometer yaw angle; determine the yaw angular velocity compensation amount according to the magnetometer data; according to the IMU angular velocity information and the The yaw angular velocity compensation amount is used to determine the corrected angular velocity; the relative value of the yaw angle is determined according to the corrected angular velocity; the fusion yaw is generated according to the initial value of the yaw angle and the relative value of the yaw angle angle.
  • FIG. 11 is a schematic diagram of an aircraft yaw angle correction device according to an embodiment of the present invention.
  • the yaw angle correction device 100 of the aircraft is applied to the aircraft, and the device includes:
  • the obtaining module 10 is used to obtain IMU data and magnetometer data, where the IMU data includes IMU acceleration information and IMU angular velocity information;
  • the determining module 20 is configured to determine the yaw angle of the magnetometer according to the magnetometer data
  • the fusion yaw angle generation module 30 is configured to generate a fusion yaw angle according to the initial value of the yaw angle and the relative value of the yaw angle.
  • FIG. 12 is a schematic structural diagram of the determining module in FIG. 11;
  • the determination module 20 includes a calibration and coordinate system conversion module 21, a stationary state detection module 22, a yaw angle deviation determination and processing module 23, and a feedback control module 24;
  • the calibration and coordinate system conversion module 21 is used for:
  • the magnetometer data of the standard magnetic field of the aircraft at the current position is compared to calculate the magnetometer yaw angle.
  • the determination module 20 further includes a static state detection module 22, and the static state detection module 22 is configured to:
  • the yaw angle deviation judgment and processing module 23 is used for:
  • the yaw angle deviation judgment and processing module is used for:
  • the yaw angle deviation judgment and processing module 23 is used for:
  • the relative deviation angle of the yaw angle is determined according to the deviation angle of the yaw angle, the ground height of the aircraft, and the flying height of the aircraft.
  • FIG. 13 is a schematic diagram of the structure of the yaw angle deviation judgment and processing module in FIG. 12;
  • the yaw angle deviation judgment and processing module 23 includes: a logical OR operation module 231 and a logical AND operation module 232;
  • the logical OR operation module 231 is used for:
  • the altitude of the aircraft meets the first preset condition
  • the derivative of the deviation angle of the yaw angle satisfies the third preset condition.
  • the first preset condition is: the height of the aircraft above the ground is greater than 0.4m, and the duration is not less than 0.5s.
  • the second preset condition is: the flying height of the aircraft is greater than 0.4m, and the duration is not less than 0.5s.
  • the third preset condition is: the absolute value of the differential angle of the yaw angle is less than 0.1 and the duration is not less than 0.5s.
  • the logical AND operation module 232 is used for:
  • the relative compensation value of the yaw angle error is determined to be the relative deviation angle of the yaw angle:
  • the relative compensation value of the yaw angle error satisfies the fourth preset condition, wherein the relative compensation value of the yaw angle error is the difference between the yaw angle deviation angle and the relative offset of the yaw angle deviation.
  • the fourth preset condition is: the absolute value of the relative compensation value of the yaw angle error is less than 0.1 and the duration is not less than 0.5s.
  • the feedback control module 24 is used to:
  • a feedback control algorithm is used to calculate the relative deviation angle of the yaw angle to determine the yaw angular velocity compensation amount.
  • an aircraft yaw angle correction device which is applied to the aircraft, and the device includes: an acquisition module for acquiring IMU data and magnetometer data, wherein the IMU data includes IMU acceleration Information and IMU angular velocity information; a determining module for: determining the magnetometer yaw angle according to the magnetometer data; determining the initial value of the yaw angle according to the magnetometer yaw angle; determining the initial value of the yaw angle according to the magnetometer data Yaw angular velocity compensation amount; determine the corrected angular velocity based on the IMU angular velocity information and the yaw angular velocity compensation amount; and determine the relative value of the yaw angle based on the corrected angular velocity; generate the fusion yaw angle
  • the module is used to generate a fusion yaw angle according to the initial value of the yaw angle and the relative value of the yaw angle.
  • FIG. 14 is a schematic diagram of the hardware structure of an aircraft according to an embodiment of the present invention.
  • the aircraft may be an unmanned aerial vehicle (UAV), an unmanned aerial vehicle or other electronic equipment.
  • UAV unmanned aerial vehicle
  • UAV unmanned aerial vehicle
  • the aircraft 1400 includes one or more processors 1401 and a memory 1402. Among them, one processor 1401 is taken as an example in FIG. 14.
  • the processor 1401 and the memory 1402 may be connected by a bus or in other ways. In FIG. 14, the connection by a bus is taken as an example.
  • the memory 1402 as a non-volatile computer-readable storage medium, can be used to store non-volatile software programs, non-volatile computer-executable programs and modules, such as the yaw angle of an aircraft in an embodiment of the present invention
  • the unit corresponding to the correction method (for example, each module or unit described in FIG. 11 to FIG. 13).
  • the processor 1401 executes various functional applications and data processing of the yaw angle correction method of the aircraft by running the non-volatile software programs, instructions, and modules stored in the memory 1402, that is, realizes the yaw of the aircraft in the above method embodiment.
  • the memory 1402 may include a high-speed random access memory, and may also include a non-volatile memory, such as at least one magnetic disk storage device, a flash memory device, or other non-volatile solid-state storage devices.
  • the memory 1402 may optionally include a memory remotely provided with respect to the processor 1401, and these remote memories may be connected to the processor 1401 through a network. Examples of the aforementioned networks include, but are not limited to, the Internet, corporate intranets, local area networks, mobile communication networks, and combinations thereof.
  • the module is stored in the memory 1402, and when executed by the one or more processors 1401, executes the aircraft yaw angle correction method in any of the foregoing method embodiments, for example, executes the above-described FIGS. 4 to 4
  • the steps shown in Fig. 10; the functions of the various modules or units described in Figs. 11 to 13 can also be realized.
  • the aircraft 1400 further includes a power system 1403, the power system 1403 is used for the aircraft to provide flight power, and the power system 1403 is connected to the processor 1401.
  • the power system 1403 includes a driving motor 14031 and an ESC 14032.
  • the ESC 14032 is electrically connected to the driving motor 14031 and used to control the driving motor 14031. Specifically, the ESC 14032 generates a control command based on the fused yaw angle obtained after the processor 1401 executes the yaw angle correction method of the aircraft, and controls the driving motor 14031 through the control command.
  • the aircraft 1400 can execute the method for correcting the yaw angle of the aircraft provided in the first embodiment of the present invention, and has functional modules and beneficial effects corresponding to the execution method.
  • the method for correcting the yaw angle of the aircraft provided in the first embodiment of the present invention please refer to the method for correcting the yaw angle of the aircraft provided in the first embodiment of the present invention.
  • the embodiment of the present invention provides a computer program product, the computer program product includes a computer program stored on a non-volatile computer-readable storage medium, the computer program includes program instructions, when the program instructions are executed by a computer At this time, the computer is made to execute the method for correcting the yaw angle of the aircraft as described above. For example, steps S10 to S70 of the method in FIG. 4 described above are executed.
  • the embodiment of the present invention also provides a non-volatile computer storage medium, the computer storage medium stores computer-executable instructions, and the computer-executable instructions are executed by one or more processors, such as a process in FIG. 14
  • the device 1401 can enable the one or more processors to execute the method for correcting the yaw angle of the aircraft in any of the foregoing method embodiments, for example, to execute the method for correcting the yaw angle of the aircraft in any of the foregoing method embodiments, for example, to execute
  • the steps shown in FIGS. 4 to 10 described above can also realize the functions of the modules or units described in FIGS. 11 to 13.
  • the above-described device or device embodiments are only illustrative.
  • the unit modules described as separate components may or may not be physically separated, and the components displayed as modular units may or may not be physical units. , Which can be located in one place, or can be distributed to multiple network module units. Some or all of the modules can be selected according to actual needs to achieve the objectives of the solutions of the embodiments.
  • each implementation manner can be implemented by means of software plus a general hardware platform, and of course, it can also be implemented by hardware.
  • the above technical solution essentially or the part that contributes to the related technology can be embodied in the form of a software product, and the computer software product can be stored in a computer-readable storage medium, such as ROM/RAM, magnetic disk , CD-ROM, etc., including several instructions until a computer device (which can be a personal computer, a server, or a network device, etc.) executes the methods described in each embodiment or some parts of the embodiment.

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Abstract

一种飞行器（10）的偏航角修正方法、装置及飞行器（10），方法包括：获取IMU数据以及磁力计数据，其中， IMU数据包括IMU加速度信息和IMU角速度信息（S10）；根据磁力计数据，确定磁力计偏航角（S20）；根据磁力计偏航角，确定偏航角初始值（S30）；根据磁力计数据，确定偏航角角速度补偿量（S40）；根据IMU角速度信息以及偏航角角速度补偿量，确定修正后的角速度（S50）；根据修正后的角速度，确定偏航角相对值（S60）；根据偏航角初始值以及偏航角相对值，生成融合偏航角（S70）。解决了室内的飞行器（10）依赖视觉信息进行偏航角修正以及室内磁干扰影响偏航角修正的问题，提高飞行器（10）在室内飞行或悬停的稳定性。

Description

一种飞行器的偏航角修正方法、装置及飞行器
本申请要求于2019年8月30日提交中国专利局、申请号为201910814909.3、申请名称为“一种飞行器的偏航角修正方法、装置及飞行器”的中国专利申请的优先权，其全部内容通过引用结合在本申请中。
技术领域
本申请涉及飞行器技术领域，特别是涉及一种飞行器的偏航角修正方法、装置及飞行器。
背景技术
飞行器，如无人飞行器(Unmanned Aerial Vehicle，UAV)，也称无人机，以其具有体积小、重量轻、机动灵活、反应快速、无人驾驶、操作要求低等优点，得到了越来越广泛的应用。无人飞行器的各个动作(或姿态)通常是通过控制无人飞行器的动力系统中的多个驱动电机不同转速实现的。其中，偏航角是对无人飞行器的飞行姿态进行控制中的重要参数，也即无人飞行器的偏航角融合对无人飞行器的姿态控制尤其重要，若无人飞行器的偏航角融合误差大，或者融合精度低，轻则无人飞行器无法按照预设的方向或轨迹飞行，重则出现刷锅现象，甚至可能失稳以致炸机。
室内环境下飞行器的磁干扰严重，并且GPS信息较差，飞行器在室内进行飞行或悬停，由于没有GPS信息修正，磁力计也受到严重干扰，没有可用的信息来进行偏航角修正，而陀螺仪积分本身存在漂移特性，因此室内飞行或悬停时，飞机容易偏航角漂移。
目前，飞行器在室内的飞行主要靠视觉信息修正或磁力计修正来修正偏航角，而视觉信息修正对于无视觉的飞机来说不可取，并且，由于视觉运算量大，对于视觉单元运算力较弱的飞机，会影响其他视觉信息的解算，而若要不影响，则需要更换更好的视觉模块，增加成本，而采用磁力计修正的方法容易受到在干扰时，飞行器偏航角偏差严重或漂移。因此，飞行器如何在室内对偏航角进行修正，是本发明需要解决的问题。
发明内容
本发明实施例提供一种飞行器的偏航角修正方法、装置及飞行器，解决室内的飞行器依赖视觉信息进行偏航角修正以及室内磁干扰影响偏航角修正的问题，提高飞行器在室内飞行或悬停的稳定性。
为解决上述技术问题，本发明实施例提供以下技术方案：
第一方面，本发明实施例提供一种飞行器的偏航角修正方法，应用于飞行器，所述方法包括：
获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；
根据所述磁力计数据，确定磁力计偏航角；
根据所述磁力计偏航角，确定偏航角初始值；
根据所述磁力计数据，确定偏航角角速度补偿量；
根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；
根据所述修正后的角速度，确定偏航角相对值；
根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。
在一些实施例中，所述根据所述磁力计数据，确定所述磁力计偏航角，包括：
对所述磁力计数据进行校准，生成校准后的磁力计数据；
获取所述飞行器的姿态角并根据所述飞行器的姿态角，生成旋转变换矩阵；
利用所述旋转变换矩阵对所述校准后的磁力计数据进行坐标变换，以生成所述地面坐标系下的磁力计数据；
根据所述地面坐标系下的磁力计数据，对比所述飞行器在当前位置下的标准磁场的磁力计数据，计算所述磁力计偏航角。
在一些实施例中，所述根据所述磁力计偏航角，确定所述偏航角初始值，包括：
判断所述飞行器在当前时刻是否由静止状态变为运动状态；
若是，则将所述磁力计偏航角作为所述偏航角初始值。
在一些实施例中，所述根据所述磁力计数据，确定偏航角角速度补偿量，包括：
根据所述磁力计偏航角，确定偏航角偏差角；
根据所述偏航角偏差角，确定偏航角相对偏差角；
根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量。
在一些实施例中，所述根据所述磁力计偏航角，确定偏航角偏差角，包括：
根据所述磁力计偏航角和上一时刻的融合偏航角，确定所述偏航角偏差角。
在一些实施例中，所述根据所述偏航角偏差角，确定偏航角相对偏差角，包括：
获取所述飞行器的对地高度和所述飞行器的飞行高度；
根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角。
在一些实施例中，所述根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角，包括：
根据所述飞行器的对地高度和所述飞行器的飞行高度，确定偏航角偏差相 对偏移量；
根据所述偏航角偏差角和所述偏航角偏差相对偏移量，确定所述偏航角相对偏差角。
在一些实施例中，所述根据所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角偏差相对偏移量，包括：
当满足以下条件中的任一项时，确定所述偏航角偏差角为所述偏航角偏差相对偏移量：
所述飞行器的对地高度满足第一预设条件；
所述飞行器的飞行高度满足第二预设条件时；以及
所述偏航角偏差角的微分满足第三预设条件。
在一些实施例中，所述第一预设条件为：
所述飞行器的对地高度大于0.4m，且持续时间不小于0.5s。
在一些实施例中，所述第二预设条件为：
所述飞行器的飞行高度大于0.4m，且持续时间不小于0.5s。
在一些实施例中，所述第三预设条件为：
所述偏航角偏差角的微分的绝对值小于0.1且持续时间不小于0.5s。
在一些实施例中，所述根据所述偏航角偏差角和所述偏航角偏差相对偏移量，确定所述偏航角相对偏差角，包括：
当均满足以下条件时，确定所述偏航角误差相对补偿值为所述偏航角相对偏差角：
所述飞行器的对地高度满足所述第一预设条件或当所述飞行器的飞行高度满足所述第二预设条件；
所述偏航角偏差角的微分满足所述第三预设条件；且
偏航角误差相对补偿值满足第四预设条件，其中，所述偏航角误差相对补偿值为所述偏航角偏差角与所述偏航角偏差相对偏移量的差。
在一些实施例中，所述第四预设条件为：
所述偏航角误差相对补偿值的绝对值小于0.1且持续时间不小于0.5s。
在一些实施例中，所述根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量，包括：
通过反馈控制算法对所述偏航角相对偏差角进行计算，以确定所述偏航角角速度补偿量。
第二方面，本发明实施例提供一种飞行器的偏航角修正装置，应用于飞行器，所述装置包括：
获取模块，用于获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；
确定模块，用于：
根据所述磁力计数据，确定磁力计偏航角；
根据所述磁力计偏航角，确定偏航角初始值；
根据所述磁力计数据，确定偏航角角速度补偿量；
根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；以及
根据所述修正后的角速度，确定偏航角相对值；
融合偏航角生成模块，用于根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。
在一些实施例中，所述确定模块包括校准及坐标系转换模块，所述校准及坐标系转换模块用于：
对所述磁力计数据进行校准，生成校准后的磁力计数据；
获取所述飞行器的姿态角并根据所述飞行器的姿态角，生成旋转变换矩阵；
利用所述旋转变换矩阵对所述校准后的磁力计数据进行坐标变换，以生成所述地面坐标系下的磁力计数据；
根据所述地面坐标系下的磁力计数据，对比所述飞行器在当前位置下的标准磁场的磁力计数据，计算所述磁力计偏航角。
在一些实施例中，所述确定模块还包括静止状态检测模块，所述静止状态检测模块用于：
判断所述飞行器在当前时刻是否由静止状态变为运动状态；
若是，则将所述磁力计偏航角作为所述偏航角初始值。
在一些实施例中，所述确定模块还包括偏航角偏差判定及处理模块，所述偏航角偏差判定及处理模块用于：
根据所述磁力计偏航角，确定偏航角偏差角；
根据所述偏航角偏差角，确定偏航角相对偏差角；
根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量。
在一些实施例中，所述偏航角偏差判定及处理模块用于：
根据所述磁力计偏航角和上一时刻的融合偏航角，确定所述偏航角偏差角。
在一些实施例中，所述偏航角偏差判定及处理模块用于：
获取所述飞行器的对地高度和所述飞行器的飞行高度；
根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角。
在一些实施例中，所述偏航角偏差判定及处理模块包括逻辑或运算模块，所述逻辑或运算模块用于：：
当满足以下条件中的任一项时，确定所述偏航角偏差角为所述偏航角偏差相对偏移量：
所述飞行器的对地高度满足第一预设条件；
所述飞行器的飞行高度满足第二预设条件时；以及
所述偏航角偏差角的微分满足第三预设条件。
在一些实施例中，所述第一预设条件为：
所述飞行器的对地高度大于0.4m，且持续时间不小于0.5s。
在一些实施例中，所述第二预设条件为：
所述飞行器的飞行高度大于0.4m，且持续时间不小于0.5s。
在一些实施例中，所述第三预设条件为：
所述偏航角偏差角的微分的绝对值小于0.1且持续时间不小于0.5s。
在一些实施例中，所述偏航角偏差判定及处理模块包括逻辑与运算模块，所述逻辑与运算模块用于：
当均满足以下条件时，确定所述偏航角误差相对补偿值为所述偏航角相对偏差角：
所述飞行器的对地高度满足所述第一预设条件或当所述飞行器的飞行高度满足所述第二预设条件；
所述偏航角偏差角的微分满足所述第三预设条件；且
偏航角误差相对补偿值满足第四预设条件，其中，所述偏航角误差相对补偿值为所述偏航角偏差角与所述偏航角偏差相对偏移量的差。
在一些实施例中，所述第四预设条件为：
所述偏航角误差相对补偿值的绝对值小于0.1且持续时间不小于0.5s。
在一些实施例中，所述确定模块包括反馈控制模块，所述反馈控制模块用于：
通过反馈控制算法对所述偏航角相对偏差角进行计算，以确定所述偏航角角速度补偿量。
第三方面，本发明实施例提供一种飞行器，包括：
机身；
机臂，与所述机身相连；
动力装置，设于所述机身和/或所述机臂，用于为所述飞行器提供飞行的动力；以及
飞行控制器，设于所述机身；
其中，所述飞行控制器包括：
至少一个处理器；以及，
与所述至少一个处理器通信连接的存储器；其中，
所述存储器存储有可被所述至少一个处理器执行的指令，所述指令被所述至少一个处理器执行，以使所述至少一个处理器能够执行如上所述的飞行器的偏航角修正方法。
第四方面，本发明实施例还提供了一种非易失性计算机可读存储介质，所述计算机可读存储介质存储有计算机可执行指令，所述计算机可执行指令用于使飞行器能够执行如上所述的飞行器的偏航角修正方法。
本发明可解决室内的飞行器依赖视觉信息进行偏航角修正以及室内磁干扰影响偏航角修正的问题，提高了飞行器在室内飞行或悬停的稳定性。
附图说明
一个或多个实施例通过与之对应的附图中的图片进行示例性说明，这些示例性说明并不构成对实施例的限定，附图中具有相同参考数字标号的元件表示为类似的元件，除非有特别申明，附图中的图不构成比例限制。
图1是本发明实施例提供的一种飞行器的具体结构图；
图2是本发明实施例提供的一种飞行器的偏航角修正方法的原理框图；
图3是本发明实施例提供的一种偏航角偏航角偏差判定及处理算法的示意图；
图4是本发明实施例提供的一种飞行器的偏航角修正方法的流程示意图；
图5是图4中的步骤S20的细化流程图；
图6是图4中的步骤S30的细化流程图；
图7是图4中的步骤S40的细化流程图；
图8是图7中的步骤S42的细化流程图；
图9是图8中的步骤S421的细化流程图；
图10是图8中的步骤S423的细化流程图；
图11是本发明实施例提供的一种飞行器的偏航角修正装置的结构示意图；
图12是图11中的确定模块的结构示意图；
图13是图12中的偏航角偏差判定及处理模块的结构示意图；
图14是本发明实施例提供的一种飞行器的硬件结构示意图；
图15是本发明实施例提供的一种飞行器的连接框图；
图16是图15中的动力系统的示意图。
具体实施方式
为使本发明实施例的目的、技术方案和优点更加清楚，下面将结合本发明实施例中的附图，对本发明实施例中的技术方案进行清楚、完整地描述，显然，所描述的实施例是本发明一部分实施例，而不是全部的实施例。基于本发明中的实施例，本领域普通技术人员在没有做出创造性劳动前提下所获得的所有其他实施例，都属于本发明保护的范围。
此外，下面所描述的本发明各个实施方式中所涉及到的技术特征只要彼此之间未构成冲突就可以相互组合。
本发明实施例提供的飞行器的偏航角修正方法可以应用到各种通过电机或马达驱动的可移动物体上，包括但不限于飞行器、机器人等。其中飞行器可包括无人飞行器(unmanned aerial vehicle，UAV)，无人飞船等。
其中，本发明实施例的飞行器的偏航角修正方法，应用于飞行器的飞行控制器。
请参阅图1，图1是本发明实施例提供的一种飞行器的具体结构图；
如图1所示，该飞行器10包括：机身11、与所述机身11相连的机臂12、设置于所述机臂12的动力装置13，连接至该机身11底部的云台14，安装在云台14上的摄像头15以及设置于机身11内的飞行控制器(图未示)。
其中，飞行控制器与动力装置13连接，动力装置13安装在所述机身11上，用于为所述飞行器10提供飞行动力。具体的，飞行控制器用于执行上述的飞行器的偏航角修正方法以修正飞行器的偏航角，并根据融合后的飞行器的偏航角生成控制指令，并将该控制指令发送给动力装置13的电调，电调通过该控制指令控制动力装置13的驱动电机。或者，飞行控制器用于执行飞行器的偏航角修正方法以修正飞行器的偏航角，并将修正后的飞行器的偏航角发送至电调，电调根据修正后的飞行器的偏航角生成控制指令，并通过该控制指令控制动力装置13的驱动电机。
机身11包括：中心壳体以及与中心壳体连接的一个或多个机臂，一个或多个机臂呈辐射状从中心壳体延伸出。机臂与中心壳体的连接可以是一体连接或者固定连接。动力装置安装于机臂上。
飞行控制器用于执行上述飞行器的偏航角修正方法以修正飞行器的偏航角，并根据修正后的飞行器的偏航角生成控制指令，并将该控制指令发送给动力装置的电调，以便电调通过该控制指令控制动力装置的驱动电机。控制器为具有一定逻辑处理能力的器件，如控制芯片、单片机、微控制单元(Microcontroller Unit，MCU)等。
动力装置13包括：电调，驱动电机和螺旋桨。电调位于机臂或中心壳体所形成的空腔内。电调分别与控制器及驱动电机连接。具体的，电调与驱动电机电连接，用于控制所述驱动电机。驱动电机安装在机臂上，驱动电机的转动轴连接螺旋桨。螺旋桨在驱动电机的驱动下产生使得飞行器10移动的力，例如，使得飞行器10移动的升力或者推力。
飞行器10完成各个规定速度、动作(或姿态)是通过电调控制驱动电机以实现的。电调全称电子调速器，根据控制信号调节飞行器10的驱动电机的转速。其中，控制器为执行上述飞行器的偏航角修正方法的执行主体，电调基于融合后的飞行器的偏航角所生成控制指令来控制驱动电机。电调控制驱动电机的原理大致为：驱动电机是将电脉冲信号转变为角位移或线位移的开环控制元件。在非超载的情况下，驱动电机的转速、停止的位置只取决于脉冲信号的频率和脉冲数，而不受负载变化的影响，当驱动器接收到一个脉冲信号，它就驱动动力装置的驱动电机按设定的方向转动一个固定的角度，它的旋转是以固定的角度运行的。因此，电调可以通过控制脉冲个数来控制角位移量，从而达到准确定位的目的；同时可以通过控制脉冲频率来控制驱动电机转动的速度和加速度，从而达到调速的目的。
目前飞行器10主要功能为航拍、影像实时传输、高危地区探测等。为了实现航拍、影像实时传输、高危地区探测等功能，飞行器10上会连接有摄像 组件。具体的，飞行器10和摄像组件通过连接结构，如减振球等进行连接。该摄像组件用于在飞行器10进行航拍的过程中，获取拍摄画面。
具体的，摄像组件包括：云台及拍摄装置。云台与飞行器10连接。其中，拍摄装置搭载于所述云台上，拍摄装置可以为图像采集装置，用于采集图像，该拍摄装置包括但不限于：相机、摄影机、摄像头、扫描仪、拍照手机等。云台用于搭载拍摄装置，以实现拍摄装置的固定、或随意调节拍摄装置的姿态(例如，改变拍摄装置的高度、倾角和/或方向)以及使所述拍摄装置稳定保持在设定的姿态上。例如，当飞行器10进行航拍时，云台主要用于使所述拍摄装置稳定保持在设定的姿态上，防止拍摄装置拍摄画面抖动，保证拍摄画面的稳定。
云台14与飞行控制器连接，以实现云台14与飞行控制器之间的数据交互。例如，飞行控制器发送偏航指令至云台14，云台14获取偏航的速度和方向指令并执行，且将执行偏航指令后所产生的数据信息发送至飞行控制器，以便飞行控制器检测当前偏航状况。
云台包括：云台电机及云台基座。其中，云台电机安装于云台基座。飞行控制器也可通过动力装置13的电调来控制云台电机，具体的，飞行控制器与电调连接，电调与云台电机电连接，飞行控制器生成云台电机控制指令，电调通过云台电机控制指令以控制云台电机。
云台基座与飞行器的机身连接，用于将摄像组件固定安装于飞行器的机身上。
云台电机分别与云台基座及拍摄装置连接。该云台可以为多轴云台，与之适应的，云台电机为多个，也即每个轴设置有一个云台电机。云台电机一方面可带动拍摄装置的转动，从而满足拍摄转轴的水平旋转和俯仰角度的调节，通过手动远程控制云台电机旋转或利用程序让电机自动旋转，从而达到全方位扫描监控的作用；另一方面，在飞行器进行航拍的过程中，通过云台电机的转动实时抵消拍摄装置受到的扰动，防止拍摄装置抖动，保证拍摄画面的稳定。
拍摄装置搭载于云台上，拍摄装置上设置有惯性测量单元(Inertial measurement unit，IMU)，该惯性测量单元用于测量物体三轴姿态角(或角速率)以及加速度的装置。一般的，一个IMU内会装有三轴的陀螺仪和三个方向的加速度计，来测量物体在三维空间中的角速度和加速度，并以此解算出物体的姿态。为了提高可靠性，还可以为每个轴配备更多的传感器。一般而言IMU要安装在飞行器的重心上。
在对飞行器的姿态进行控制的过程中，飞行器的偏航角是对飞行器的姿态进行控制中的重要参数，需要基于飞行器的偏航角，来控制驱动电机。通过飞行器的控制器实时获取飞行器的偏航角，为飞行器的姿态控制提供必要的姿态信息。也即飞行器的偏航角正确估算对飞行器的姿态控制尤其重要，若飞行器的偏航角估算错误，飞行器轻则无法按照预设的方向或轨迹飞行，重则可能失稳以致炸机。
在室内环境中，由于没有GPS信息修正，磁力计也受到严重干扰，因此导致存在缺乏足够的可用信息来进行偏航角的修正的问题，而且，由于陀螺仪积分本身存在漂移特性，因此在室内飞行或悬停时，飞行器容易发生偏航角偏移。
目前，飞行器在室内的飞行主要靠视觉信息修正或磁力计修正来修正偏航角，而视觉信息修正对于无视觉的飞机来说不可取，并且，由于视觉运算量大，对于视觉单元运算力较弱的飞机，会影响其他视觉信息的解算，而若要不影响，则需要更换更好的视觉模块，增加成本，而采用磁力计修正的方法容易受到在干扰时，飞行器偏航角偏差严重或漂移。
因此，基于上述问题，本发明实施例主要目的在于提供一种飞行器的偏航角修正方法、装置及飞行器，可以基于IMU数据以及磁力计数据，对飞行器的偏航角进行修正，解决室内的飞行器依赖视觉信息进行偏航角修正以及室内磁干扰影响偏航角修正的问题，从而提高飞行器在室内飞行或悬停的稳定性。
本发明实施例通过获取IMU数据以及磁力计数据，计算偏航角初始值以及偏航角相对值，并对偏航角初始值以及偏航角相对值进行融合，该融合方法可避免室内的磁干扰，也即在缺乏GPS信号与强磁干扰环境下，也能保证飞行器的飞行或悬停的稳定性。
下面结合附图，对本发明实施例作进一步阐述。
实施例一
请参阅图2，图2是本发明实施例提供的一种飞行器的偏航角修正方法的原理框图；
如图2所示，通过获取IMU数据以及磁力计数据，通过对所述IMU数据进行校准、坐标系转换，获取IMU加速度以及IMU角速度，对飞行器进行静止检测，生成静止标志位信号，将所述静止标志位信号输入飞行器的使能模块，根据所述静止标志位信号的上升沿，将磁力计偏航角输出作为偏航角初始值，并且，通过对磁力计数据进行标准矩阵旋转变换，生成磁力计偏航角，根据磁力计偏航角与当前的融合偏航角，计算偏航角偏差角，对所述偏航角偏差角进行判定及处理，生成偏航角相对偏差角，根据所述偏航角相对偏差角，生成偏航角角速度补偿量，还通过将IMU角速度以及根据磁力计数据生成的偏航角角速度补偿量进行融合，生成修正后的加速度，对所述修正后的加速度进行积分，得到偏航角相对值，将所述偏航角初始值和偏航角相对值进行融合，生成融合偏航角。
请再参阅图3，图3是本发明实施例提供的一种偏航角偏航角偏差判定及处理算法的示意图；
如图3所示，通过获取对地高度、飞行高度等数据，并对上述数据进行相应的逻辑运算或处理，生成偏航角相对偏差角。
请参阅图4，图4是本发明实施例提供的一种飞行器的偏航角修正方法的流程示意图；
其中，该飞行器的偏航角修正方法可由各种具有一定逻辑处理能力的电子 设备执行，如飞行器、控制芯片等，该飞行器可以包括无人机、无人船等。以下电子设备以飞行器为例进行说明。其中，飞行器连接有云台，云台包括云台电机及云台基座，其中，云台可以为多轴云台，如两轴云台、三轴云台，以下三轴云台为例进行说明。对于该飞行器及云台的具体结构的描述可以参考上述描述，因此，在此处不作赘述。
如图4所示，所述方法应用于飞行器，比如，无人机，所述方法包括：
步骤S10：获取IMU数据以及磁力计数据，所述IMU数据包括IMU加速度信息和IMU角速度信息；
具体的，所述飞行器设置有姿态传感器组件，所述姿态传感器组件包括：惯性测量单元(Inertial measurement unit，IMU)、磁力计等，其中，所述IMU用于获取IMU数据，所述磁力计用于获取磁力计数据，所述惯性测量单元包括陀螺仪以及加速度计，所述陀螺仪用于获取IMU角速度，所述加速度计用于获取IMU加速度信息，所述IMU数据包括：IMU加速度信息以及IMU角速度信息，所述磁力计数据包括：磁场强度信息。
具体的，通过惯性测量单元获取IMU数据，并对所述IMU数据进行校准、坐标系转换，生成IMU加速度信息以及IMU角速度信息，其中，所述IMU加速度信息为惯性测量单元的测量数据经过校准矩阵进行校准以及机体坐标系到地面坐标系的坐标变换之后，所得到的地面坐标系下的加速度信息。可以理解的是，所述校准矩阵是用户在要飞行的地方校准得到的，校准矩阵在地球上任意地方都不同，飞行器在报磁力计干扰，要求用户校准之后才能确定所述校准矩阵。
其中，所述机体坐标系到地面坐标系的转换通过旋转变换矩阵完成，具体的，根据所述飞行器的姿态角，生成旋转变换矩阵，通过所述旋转变换矩阵，将所述IMU数据从机体坐标系转换到地面坐标系，生成所述IMU加速度信息以及所述IMU角速度信息。具体的，所述飞行器的姿态角包括：偏航角、俯仰角以及翻滚角，其中，所述偏航角为当前的融合偏航角，即实时的融合偏航角会用于计算旋转变换矩阵，进而用于下一次的融合，不断更新所述融合偏航角。例如：所述旋转变换矩阵为3*3的矩阵，其中包含了所述偏航角、俯仰角、翻滚角的正弦余弦函数，并根据具体情况选择不同的函数，一般而言，通过先转动偏航角，再转动俯仰角，最后转动翻滚角，例如：所述旋转变换矩阵为：
Figure PCTCN2020111310-appb-000001
其中，(φ,θ,ψ)为所述姿态角，φ为所述姿态角中的翻滚角，θ为所述姿态角中的俯仰角，ψ为所述姿态角中的偏航角。
步骤S20：根据所述磁力计数据，确定磁力计偏航角；
其中，所述磁力计数据通过磁力计获取，所述磁力计数据包括：磁场强度信息，所述磁场强度为三轴磁场强度，由于磁力计测量的磁力计数据是机体坐 标系下的三轴磁场强度，因此需要通过校准矩阵去除bias和交叉耦合，并且，通过旋转矩阵将其变换至地面坐标系下。具体的，请再参阅图5，图5是图4中的步骤S20的细化流程图；
如图5所示，所述根据所述磁力计数据，确定磁力计偏航角，包括：
步骤S21：对所述磁力计数据进行校准，生成校准后的磁力计数据；
其中，根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；具体的，所述预设的校准矩阵是用户在要飞行的地方校准得到的，校准矩阵在地球上任意地方都不同，飞行器在报磁力计干扰，要求用户校准之后才能确定所述校准矩阵。
步骤S22：获取所述飞行器的姿态角并根据所述飞行器的姿态角，生成旋转变换矩阵；
具体的，所述旋转变换矩阵用于将机体坐标系转换为地面坐标系，所述飞行器的姿态角包括：偏航角、俯仰角、翻滚角，通过获取所述飞行器的姿态角，其中，所述偏航角为当前的融合偏航角，即实时的融合偏航角会用于计算旋转变换矩阵，进而用于下一次的融合，不断更新所述融合偏航角。所述旋转变换矩阵为3*3的矩阵，其中包含了所述偏航角、俯仰角、翻滚角的正弦余弦函数，并根据具体情况选择不同的函数，一般而言，通过先转动偏航角，再转动俯仰角，最后转动翻滚角，例如：所述旋转变换矩阵为：
Figure PCTCN2020111310-appb-000002
其中，(φ,θ,ψ)为所述姿态角，φ为所述姿态角中的翻滚角，θ为所述姿态角中的俯仰角，ψ为所述姿态角中的偏航角。
步骤S23：利用所述旋转变换矩阵对所述校准后的磁力计数据进行坐标变换，以生成所述地面坐标系下的磁力计数据；
具体的，将所述校准后的磁力计数据，即所述磁场强度，乘以所述旋转变换矩阵，生成所述地面坐标系下的磁场强度，相当于对所述磁力计数据进行机体坐标系和地面坐标系之间的变换，通过坐标变换后，以生成所述地面坐标系下的磁力计数据。
步骤S24：根据所述地面坐标系下的磁力计数据，对比所述飞行器在当前位置下的标准磁场的磁力计数据，计算所述磁力计偏航角。
具体的，所述飞行器的当前位置对应一标准磁场，所述磁力计三轴读数组成一个向量，所述当前位置的标准磁场的磁力计数据对应一向量，通过将所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算两者之间的向量夹角，将所述向量夹角作为所述磁力计偏航角。
步骤S30：根据所述磁力计偏航角，确定偏航角初始值；
具体的，所述飞行器设置有使能模块，所述使能模块包括输入端和输出端，当所述使能模块接受到静止标志位信号的上升沿时，所述使能模块将所述磁力 计偏航角输出，作为偏航角初始值。
具体的，请再参阅图6，图6是图4中的步骤S30的细化流程图；
如图6所示，所述根据所述磁力计偏航角，确定偏航角初始值，包括：
步骤S31：判断所述飞行器在当前时刻是否由静止状态变为运动状态；
具体的，根据所述飞行器的静止状态，确定所述飞行器的静止标志位的信号；其中，所述飞行器的静止标志位用于表征所述飞行器的静止状态，所述根据所述飞行器的静止状态，确定所述飞行器的静止标志位的信号，包括：获取所述IMU数据中的IMU加速度以及IMU角速度，通过对所述IMU加速度以及IMU角速度进行静止检测，确定所述飞行器的静止状态，若所述飞行器的静止状态为静止，则所述静止状态位的值为1，若所述飞行器的静止状态为运动，则所述静止状态位的值为0，可以理解的是，静止时所述IMU加速度以及IMU角速度的数据波动特别小，很稳定，运动时所述IMU加速度以及IMU角速度的数据波动变化大，根据此原理确定所述飞行器的静止状态。其中，所述方法具体包括：根据旋转变换矩阵，对所述IMU数据进行机体坐标系和地面坐标系的转换，生成地面坐标系下的IMU加速度以及地面坐标系下的IMU角速度；根据所述地面坐标系下的IMU加速度以及地面坐标系下的IMU角速度，确定所述飞行器的静止状态，生成所述飞行器的静止标志位。
步骤S32：若是，则将所述磁力计偏航角作为所述偏航角初始值。
具体的，将所述静止状态位的信号输入所述飞行器的使能模块，当所述使能模块检测到所述静止状态位的信号存在上升沿，将所述磁力计偏航角输出作为所述偏航角初始值。其中，所述静止状态位的信号在飞行器静止时为0，在飞行器运动时为1，将所述静止状态位的信号输入所述飞行器的使能模块，当所述使能模块检测到所述静止状态位的信号存在上升沿，即所述静止状态位从0到1时，所述使能模块将所述磁力计偏航角输出作为所述偏航角初始值。
步骤S40：根据所述磁力计数据，确定偏航角角速度补偿量；
具体的，由于室内磁干扰严重，因此IMU测量的IMU数据需要进行修正，所述偏航角角速度补偿量用于对IMU获取的IMU角速度进行修正，需要通过所述磁力计偏航角，确定偏航角角速度补偿量，具体的，请参阅图7，图7是图4中的步骤S40的细化流程图；
如图7所示，所述根据所述磁力计偏航角，确定偏航角角速度补偿量，包括：
步骤S41：根据所述磁力计偏航角，确定偏航角偏差角；
具体的，所述根据所述磁力计偏航角，确定偏航角偏差角，包括：根据所述磁力计偏航角和上一时刻的融合偏航角，确定所述偏航角偏差角。其中，所述当前的融合偏航角为实时的融合偏航角，将所述当前的融合偏航角与所述磁力计偏航角的差值作为所述偏航角偏差角。
步骤S42：根据所述偏航角偏差角，确定偏航角相对偏差角；
具体的，所述根据所述偏航角偏差角，确定偏航角相对偏差角，包括：
获取所述飞行器的对地高度和所述飞行器的飞行高度；
根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角。
具体的，所述根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角，包括：
根据所述飞行器的对地高度和所述飞行器的飞行高度，确定偏航角偏差相对偏移量；
根据所述偏航角偏差角和所述偏航角偏差相对偏移量，确定所述偏航角相对偏差角。
具体的，所述根据所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角偏差相对偏移量，包括：
当满足以下条件中的任一项时，确定所述偏航角偏差角为所述偏航角偏差相对偏移量：
所述飞行器的对地高度满足第一预设条件；
所述飞行器的飞行高度满足第二预设条件时；以及
所述偏航角偏差角的微分满足第三预设条件。
其中，所述第一预设条件为：所述飞行器的对地高度大于0.4m，且持续时间不小于0.5s，所述第二预设条件为：所述飞行器的飞行高度大于0.4m，且持续时间不小于0.5s，所述第三预设条件为：所述偏航角偏差角的微分的绝对值小于0.1且持续时间不小于0.5s。
其中，所述偏航角相对偏差角为对所述偏航角偏差角进行判定及处理后的偏航角相对偏差量。具体的，请再参阅图8，图8是图7中的步骤S42的细化流程图；
如图8所示，所述根据所述偏航角偏差角，确定偏航角相对偏差角，包括：
步骤S421：根据所述偏航角偏差角，确定偏航角偏差相对偏移量；
具体的，所述飞行器设置有锁存模块，所述锁存模块用于锁存所述偏航角偏差角，当接收到有效信号时，所述锁存模块将所述偏航角偏差角输出作为所述偏航角偏差相对偏移量，并持续锁存所述偏航角偏差角，直至下一个有效信号到来，才更新所述偏航角偏差相对偏移量。
请再参阅图9，图9是图8中的步骤S421的细化流程图；
如图9所示，所述根据所述偏航角偏差角，确定偏航角偏差相对偏移量，包括：
步骤S4211：获取高度重置脉冲信号以及微分重置脉冲信号；
具体的，所述获取高度重置脉冲，包括：
获取所述飞行器的对地高度以及飞行高度；
具体的，所述飞行器的对地高度指的是飞行器飞行时所在位置距离当前正下方地面的距离，所述飞行器的飞行高度指的是所述飞行器飞行时所在的位置距离起飞点的高度落差，所述飞行器设置有超声波传感器和/或TOF传感器 (Time of Flight，TOF)，所述飞行器的对地高度和飞行高度通过所述超声波传感器和/或TOF传感器获取。
根据所述对地高度以及飞行高度，确定所述飞行器的高度判定标志位信号；
其中，所述高度判定标志位信号用于避免地板里面钢筋混凝土的磁干扰，具体的，预设对地高度阈值及其持续时间阈值，若所述飞行器的对地高度大于所述对地高度阈值，并且所述对地高度大于所述对地高度阈值的持续时间大于所述持续时间阈值，则所述高度判定标志位信号输出高电平，否则，所述高度判定标志位信号输出低电平；或者，预设飞行高度阈值及其持续时间阈值，若所述飞行器的飞行高度大于所述飞行高度阈值，并且所述飞行器的飞行高度大于所述飞行高度阈值的持续时间大于所述持续时间阈值，则所述高度判定标志位信号输出高电平，否则，所述高度判定标志位输出低电平，例如：所述预设对地高度阈值及其持续时间阈值分别为0.4m和0.5s，预设飞行高度阈值及其持续时间阈值分别为0.4m和0.5s，当飞机对地高度>0.4m并持续0.5s以上时，高度判定标志位信号输出1；或者，当飞机飞行高度>0.4m并持续0.5s以上时，高度判定标志位信号输出1；否则，高度判定标志位信号输出0。可以理解的是，本发明实施例中的对地高度阈值及其持续时间阈值，以及，飞行高度阈值及其持续时间阈值可以根据具体需求具体设置，均在本发明的保护范围内。
若所述飞行器的高度判定标志位的信号存在上升沿，则生成高度重置脉冲信号。
具体的，所述飞行器的控制器设置有上升沿检测模块，所述上升沿检测模块获取所述高度判定标志位信号，并判断所述高度判定标志位信号是否存在上升沿，若是，则生成高度重置脉冲信号。
具体的，所述获取微分重置脉冲信号，包括：
对所述偏航角偏差角进行微分以及滤波处理，生成偏航角微分值；
具体的，对所述偏航角偏差角进行微分及滤波处理，通过对所述偏航角进行微分，获取初始的偏航角微分值，对所述初始的偏航角微分值进行滤波处理，生成最终的偏航角微分值，即所述偏航角微分值，其中，所述滤波处理包括低通滤波和卡尔曼滤波，分别通过所述飞行器的低通滤波器和卡尔曼滤波器完成。
对所述偏航角微分值进行逻辑判断，确定偏航角微分判定标志位信号；
具体的，所述飞行器的控制器设置有判断逻辑模块，所述判断逻辑模块获取所述偏航角微分值，对所述偏航角微分值进行逻辑判断，生成所述偏航角微分判定标志位信号，其中，所述判断逻辑模块对所述偏航角微分值的判断逻辑，包括：预设微分值绝对值阈值及其持续时间阈值，若所述偏航角微分值小于所述微分值绝对值阈值，并且，所述偏航角微分值小于所述微分值绝对值阈值的持续时间大于所述持续时间阈值，则所述偏航角微分判定标志位信号输出高电平，否则，所述偏航角微分判定标志位信号输出低电平，例如：预设所述微分 值绝对值阈值为0.1，所述持续时间阈值为0.5s，若所述偏航角微分值的绝对值小于0.1，并且，所述偏航角微分值的绝对值小于0.1的持续时间大于所述0.5s，则所述偏航角微分判定标志位信号输出1，否则，所述偏航角微分判定标志位信号输出0。可以理解的是，所述微分值绝对值阈值及其持续时间阈值可以根据具体需求具体设置，均在本发明的保护范围内。
若所述偏航角微分判定标志位的信号存在上升沿，则生成微分重置脉冲信号。
具体的，所述飞行器的上升沿检测模块获取所述偏航角微分判定标志位信号，并判断所述偏航角微分判定标志位信号是否存在上升沿，若存在，则生成微分重置脉冲信号，并且所述微分重置脉冲信号输出高电平，否则，所述微分重置脉冲信号输出低电平。
步骤S4212：对所述高度重置脉冲信号以及所述微分重置脉冲信号进行逻辑判断，生成偏航角相对偏移量重置脉冲标志位信号；
具体的，对所述高度重置脉冲信号以及所述微分重置脉冲信号进行逻辑或运算，并将逻辑或运算之后的结果作为所述偏航角相对偏移量重置脉冲标志位信号。例如：所述高度重置脉冲信号或者所述微分重置脉冲信号任意一个为高电平，则所述偏航角相对偏移量重置脉冲标志位信号为高电平，若所述高度重置脉冲信号以及所述微分重置脉冲信号均为低电平，则所述偏航角相对偏移量重置脉冲标志位信号为低电平。
步骤S4213：向所述飞行器的锁存模块发送所述偏航角相对偏移量重置脉冲标志位信号，若所述飞行器的锁存模块接收到所述偏航角相对偏移量重置脉冲标志位的有效信号，将所述偏航角偏差角输出作为偏航角偏差相对偏移量。
具体的，所述有效信号指的是高电平信号，即1，若所述偏航角相对偏移量重置脉冲标志位信号为高电平，则所述飞行器的锁存模块将所述偏航角偏差角输出作为偏航角偏差相对偏移量，若所述偏航角相对偏移量重置脉冲标志位信号为低电平，则所述飞行器的锁存模块持续锁存所述偏航角偏差角，直至接受到所述偏航角相对偏移量重置脉冲标志位信号的有效信号，即所述偏航角相对偏移量重置脉冲标志位信号为高电平信号。
步骤S422：根据所述偏航角偏差角以及所述偏航角偏差相对偏移量，确定偏航角误差相对补偿值；
具体的，通过公式：偏航角误差相对补偿值＝偏航角偏差角-偏航角偏差相对偏移量，计算所述偏航角误差相对补偿值。
步骤S423：根据所述偏航角误差相对补偿值，确定所述偏航角相对偏差角。
具体的，请参阅图10，图10是图8中的步骤S423的细化流程图；
如图10所示，所述根据所述偏航角误差相对补偿值，确定所述偏航角相对偏差角，包括：
步骤S4231：对所述偏航角误差相对补偿值进行逻辑判断，生成偏航角误 差相对补偿值判定标志位信号；
具体的，所述飞行器的判断逻辑模块获取所述偏航角误差相对补偿值，并对所述偏航角误差相对补偿值进行逻辑判断，其中，所述判断逻辑模块对所述偏航角误差相对补偿值的判断逻辑，包括：预设补偿值绝对值阈值及其持续时间阈值，若所述偏航角误差相对补偿值小于所述补偿值绝对值阈值，并且，所述偏航角误差相对补偿值小于所述补偿值绝对值阈值的持续时间大于所述持续时间阈值，则所述偏航角误差相对补偿值判定标志位信号输出高电平，否则，所述偏航角误差相对补偿值判定标志位信号输出低电平，例如：预设所述补偿值微分值绝对值阈值为0.1，所述持续时间阈值为0.5s，若所述偏航角误差相对补偿值的绝对值小于0.1，并且，所述偏航角误差相对补偿值的绝对值小于0.1的持续时间大于所述0.5s，则所述偏航角误差相对补偿值判定标志位信号输出1，否则，所述偏航角误差相对补偿值判定标志位信号输出0。可以理解的是，所述补偿值绝对值阈值及其持续时间阈值可以根据具体需求具体设置，均在本发明的保护范围内。
步骤S4232：将所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号进行逻辑判断，生成偏航角偏差补偿标志位；
具体的，对所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号进行逻辑与运算，确定所述偏航角偏差补偿标志位的值，所述偏航角偏差补偿标志位用于确定所述偏航角误差相对补偿值能否用于补偿，若所述偏航角偏差补偿标志位为高电平，则代表可以用来补偿，若所述偏航角偏差补偿标志位为低电平，则代表不能用于补偿。例如：所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号均为高电平，则所述飞行器的判断逻辑模块对所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号进行逻辑与运算的值为高电平，则所述偏航角偏差补偿标志位输出高电平，或者，若所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号任意一个为低电平，则所述飞行器的判断逻辑模块对所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号进行逻辑与运算的值为低电平，则所述偏航角偏差补偿标志位输出低电平。
步骤S4233：将所述偏航角偏差补偿标志位的信号输入所述飞行器的使能模块，若所述飞行器的使能模块接收到所述偏航角偏差补偿标志位的有效信号，则所述使能模块将所述偏航角误差相对补偿值输出作为所述偏航角相对偏差角。
具体的，若对所述高度判定标志位信号、偏航角微分判定标志位信号以及所述偏航角误差相对补偿值判定标志位信号进行逻辑与运算的结果为高电平，即所述偏航角偏差补偿标志位输出有效信号，则所述飞行器的使能模块将所述 偏航角误差相对补偿值输出作为所述偏航角相对偏差角。
步骤S43：根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量。
具体的，所述飞行器设置有反馈控制器，将所述偏航角相对偏差角输入所述反馈控制器，所述反馈控制器通过反馈控制算法对所述偏航角相对偏差角进行计算，确定所述偏航角角速度补偿量，例如：所述偏航角角速度补偿量与所述偏航角相对偏差角负相关，例如：偏航角角速度补偿量＝-K*偏航角相对偏差角，其中，所述K为工程师设计值。
步骤S50：根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；
具体的，对所述IMU角速度信息以及所述偏航角角速度补偿量进行求和，将求得的和作为所述修正后的角速度，例如：所述修正后的角速度＝所述IMU角速度信息+所述偏航角角速度补偿量。
步骤S60：根据所述修正后的角速度，确定偏航角相对值；
具体的，对所述修正后的角速度进行积分，将所述修正后的角速度的积分值作为所述偏航角相对值。
步骤S70：根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。
具体的，将所述偏航角初始值和所述偏航角相对值进行求和，将求和的结果作为所述融合偏航角，例如：所述融合偏航角＝所述偏航角初始值+所述偏航角相对值。
具体的，所述飞行器的每一个采样步长都有做一次误差计算，通过反馈回路一直不停歇的无限进行。融合好的融合偏航角要去与磁力计的偏航角作差，生成偏航角误差角，无休止的进行，直到飞机断电。所述融合偏航角也是无休止的在更新，每一个采样时刻对应唯一的融合偏航角，所述偏航角误差角为所述磁力计偏航角和所述融合偏航角的差值。
在本发明实施例中，通过提供一种飞行器的偏航角修正方法，所述方法包括：获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；根据所述磁力计数据，确定磁力计偏航角；根据所述磁力计偏航角，确定偏航角初始值；根据所述磁力计数据，确定偏航角角速度补偿量；根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；根据所述修正后的角速度，确定偏航角相对值；根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。通过获取IMU数据以及磁力计数据，计算偏航角初始值以及偏航角相对值，并对偏航角初始值以及偏航角相对值进行融合，充分地、合理地、巧妙地利用磁力计信息来进行Yaw角修正，并且，不依赖视觉信息，只通过受干扰的磁力计来稳定偏航角，提升融合精度，可以有效避免室内的磁干扰，也即在缺乏GPS信号与强磁干扰环境下，也能保证飞行器的飞行或悬停的稳定性。
实施例二
请参阅图11，图11是本发明实施例提供的一种飞行器的偏航角修正装置的示意图；
如图11所示，该飞行器的偏航角修正装置100，应用于飞行器，所述装置包括：
获取模块10，用于获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；
确定模块20，用于根据所述磁力计数据，确定磁力计偏航角；
根据所述磁力计偏航角，确定偏航角初始值；
根据所述磁力计偏航角，确定偏航角角速度补偿量；
根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；
根据所述修正后的角速度，确定偏航角相对值；
融合偏航角生成模块30，用于根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。
请再参阅图12，图12是图11中的确定模块的结构示意图；
如图12所示，所述确定模块20包括校准及坐标系转换模块21、静止状态检测模块22、偏航角偏差判定及处理模块23以及反馈控制模块24；
具体的，所述校准及坐标系转换模块21用于：
对所述磁力计数据进行校准，生成校准后的磁力计数据；
获取所述飞行器的姿态角并根据所述飞行器的姿态角，生成旋转变换矩阵；
利用所述旋转变换矩阵对所述校准后的磁力计数据进行坐标变换，以生成所述地面坐标系下的磁力计数据；
根据所述地面坐标系下的磁力计数据，对比所述飞行器在当前位置下的标准磁场的磁力计数据，计算所述磁力计偏航角。
在本发明实施例中，所述确定模块20还包括静止状态检测模块22，所述静止状态检测模块22用于：
判断所述飞行器在当前时刻是否由静止状态变为运动状态；
若是，则将所述磁力计偏航角作为所述偏航角初始值。
具体的，所述偏航角偏差判定及处理模块23用于：
根据所述磁力计偏航角，确定偏航角偏差角；
根据所述偏航角偏差角，确定偏航角相对偏差角；
根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量。
在本发明实施例中，所述偏航角偏差判定及处理模块，用于：
根据所述磁力计偏航角和上一时刻的融合偏航角，确定所述偏航角偏差角。
具体的，所述偏航角偏差判定及处理模块23，用于：
获取所述飞行器的对地高度和所述飞行器的飞行高度；
根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角。
具体的，请再参阅图13，图13是图12中的偏航角偏差判定及处理模块的结构示意图；
如图13所示，所述偏航角偏差判定及处理模块23，包括：逻辑或运算模块231以及逻辑与运算模块232；
具体的，所述逻辑或运算模块231用于：
当满足以下条件中的任一项时，确定所述偏航角偏差角为所述偏航角偏差相对偏移量：
所述飞行器的对地高度满足第一预设条件；
所述飞行器的飞行高度满足第二预设条件时；以及
所述偏航角偏差角的微分满足第三预设条件。
其中，所述第一预设条件为：所述飞行器的对地高度大于0.4m，且持续时间不小于0.5s。
其中，所述第二预设条件为：所述飞行器的飞行高度大于0.4m，且持续时间不小于0.5s。
其中，所述第三预设条件为：所述偏航角偏差角的微分的绝对值小于0.1且持续时间不小于0.5s。
具体的，所述逻辑与运算模块232用于：
当均满足以下条件时，确定所述偏航角误差相对补偿值为所述偏航角相对偏差角：
所述飞行器的对地高度满足所述第一预设条件或当所述飞行器的飞行高度满足所述第二预设条件；
所述偏航角偏差角的微分满足所述第三预设条件；且
偏航角误差相对补偿值满足第四预设条件，其中，所述偏航角误差相对补偿值为所述偏航角偏差角与所述偏航角偏差相对偏移量的差。
其中，所述第四预设条件为：所述偏航角误差相对补偿值的绝对值小于0.1且持续时间不小于0.5s。
具体的，所述反馈控制模块24用于：
通过反馈控制算法对所述偏航角相对偏差角进行计算，以确定所述偏航角角速度补偿量。
在本发明实施例中，通过提供一种飞行器的偏航角修正装置，应用于飞行器，所述装置包括：获取模块，用于获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；确定模块，用于：根据所述磁力计数据，确定磁力计偏航角；根据所述磁力计偏航角，确定偏航角初始值；根据所述磁力计数据，确定偏航角角速度补偿量；根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；以及根据所述修正后的角速度，确定偏航角相对值；融合偏航角生成模块，用于根据所述偏航角初 始值以及所述偏航角相对值，生成融合偏航角。通过上述方式，本发明解决室内的飞行器依赖视觉信息进行偏航角修正以及室内磁干扰影响偏航角修正的问题，提高飞行器在室内飞行或悬停的稳定性。
请参阅图14，图14是本发明实施例提供一种飞行器的硬件结构示意图。其中，该飞行器可以是无人飞行器(unmanned aerial vehicle，UAV)，无人飞船等电子设备。
如图14所示，该飞行器1400包括一个或多个处理器1401以及存储器1402。其中，图14中以一个处理器1401为例。
处理器1401和存储器1402可以通过总线或者其他方式连接，图14中以通过总线连接为例。
存储器1402作为一种非易失性计算机可读存储介质，可用于存储非易失性软件程序、非易失性计算机可执行程序以及模块，如本发明实施例中的一种飞行器的偏航角修正方法对应的单元(例如，图11至图13所述的各个模块或单元)。处理器1401通过运行存储在存储器1402中的非易失性软件程序、指令以及模块，从而执行飞行器的偏航角修正方法的各种功能应用以及数据处理，即实现上述方法实施例飞行器的偏航角修正方法以及上述装置实施例的各个模块和单元的功能。
存储器1402可以包括高速随机存取存储器，还可以包括非易失性存储器，例如至少一个磁盘存储器件、闪存器件、或其他非易失性固态存储器件。在一些实施例中，存储器1402可选包括相对于处理器1401远程设置的存储器，这些远程存储器可以通过网络连接至处理器1401。上述网络的实例包括但不限于互联网、企业内部网、局域网、移动通信网及其组合。
所述模块存储在所述存储器1402中，当被所述一个或者多个处理器1401执行时，执行上述任意方法实施例中的飞行器的偏航角修正方法，例如，执行以上描述的图4至图10所示的各个步骤；也可实现图11至图13所述的各个模块或单元的功能。
请参阅图15和图16，所述飞行器1400还包括动力系统1403，所述动力系统1403用于飞行器提供飞行动力，所述动力系统1403与处理器1401连接。所述动力系统1403包括：驱动电机14031及电调14032，所述电调14032与驱动电机14031电连接，用于控制所述驱动电机14031。具体的，所述电调14032基于处理器1401执行上述飞行器的偏航角修正方法后得到的融合偏航角，生成控制指令，通过控制指令控制该驱动电机14031。
所述飞行器1400可执行本发明实施例一所提供的飞行器的偏航角修正方法，具备执行方法相应的功能模块和有益效果。未在飞行器实施例中详尽描述的技术细节，可参见本发明实施例一所提供的飞行器的偏航角修正方法。
本发明实施例提供了一种计算机程序产品，所述计算机程序产品包括存储在非易失性计算机可读存储介质上的计算机程序，所述计算机程序包括程序指令，当所述程序指令被计算机执行时，使所述计算机执行如上所述的飞行器的 偏航角修正方法。例如，执行以上描述的图4中的方法步骤S10至步骤S70。
本发明实施例还提供了一种非易失性计算机存储介质，所述计算机存储介质存储有计算机可执行指令，该计算机可执行指令被一个或多个处理器执行，例如图14中的一个处理器1401，可使得上述一个或多个处理器可执行上述任意方法实施例中的飞行器的偏航角修正方法，例如，执行上述任意方法实施例中的飞行器的偏航角修正方法，例如，执行以上描述的图4至图10所示的各个步骤；也可实现图11至图13所述的各个模块或单元的功能。
以上所描述的装置或设备实施例仅仅是示意性的，其中所述作为分离部件说明的单元模块可以是或者也可以不是物理上分开的，作为模块单元显示的部件可以是或者也可以不是物理单元，即可以位于一个地方，或者也可以分布到多个网络模块单元上。可以根据实际的需要选择其中的部分或者全部模块来实现本实施例方案的目的。
通过以上的实施方式的描述，本领域的技术人员可以清楚地了解到各实施方式可借助软件加通用硬件平台的方式来实现，当然也可以通过硬件。基于这样的理解，上述技术方案本质上或者说对相关技术做出贡献的部分可以以软件产品的形式体现出来，该计算机软件产品可以存储在计算机可读存储介质中，如ROM/RAM、磁碟、光盘等，包括若干指令用直至得一台计算机设备(可以是个人计算机，服务器，或者网络设备等)执行各个实施例或者实施例的某些部分所述的方法。
最后应说明的是：以上实施例仅用以说明本发明的技术方案，而非对其限制；在本发明的思路下，以上实施例或者不同实施例中的技术特征之间也可以进行组合，步骤可以以任意顺序实现，并存在如上所述的本发明的不同方面的许多其它变化，为了简明，它们没有在细节中提供；尽管参照前述实施例对本发明进行了详细的说明，本领域的普通技术人员应当理解：其依然可以对前述各实施例所记载的技术方案进行修改，或者对其中部分技术特征进行等同替换；而这些修改或者替换，并不使相应技术方案的本质脱离本申请各实施例技术方案的范围。

Claims (28)

  1. 一种飞行器的偏航角修正方法，应用于飞行器，其特征在于，所述方法包括：
    获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；
    根据所述磁力计数据，确定磁力计偏航角；
    根据所述磁力计偏航角，确定偏航角初始值；
    根据所述磁力计数据，确定偏航角角速度补偿量；
    根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；
    根据所述修正后的角速度，确定偏航角相对值；
    根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。
  2. 根据权利要求1所述的方法，其特征在于，所述根据所述磁力计数据，确定所述磁力计偏航角，包括：
    对所述磁力计数据进行校准，生成校准后的磁力计数据；
    获取所述飞行器的姿态角并根据所述飞行器的姿态角，生成旋转变换矩阵；
    利用所述旋转变换矩阵对所述校准后的磁力计数据进行坐标变换，以生成所述地面坐标系下的磁力计数据；
    根据所述地面坐标系下的磁力计数据，对比所述飞行器在当前位置下的标准磁场的磁力计数据，计算所述磁力计偏航角。
  3. 根据权利要求1所述的方法，其特征在于，所述根据所述磁力计偏航角，确定所述偏航角初始值，包括：
    判断所述飞行器在当前时刻是否由静止状态变为运动状态；
    若是，则将所述磁力计偏航角作为所述偏航角初始值。
  4. 根据权利要求1所述的方法，其特征在于，所述根据所述磁力计数据，确定偏航角角速度补偿量，包括：
    根据所述磁力计偏航角，确定偏航角偏差角；
    根据所述偏航角偏差角，确定偏航角相对偏差角；
    根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量。
  5. 根据权利要求4所述的方法，其特征在于，所述根据所述磁力计偏航角，确定偏航角偏差角，包括：
    根据所述磁力计偏航角和上一时刻的融合偏航角，确定所述偏航角偏差 角。
  6. 根据权利要求4所述的方法，其特征在于，所述根据所述偏航角偏差角，确定偏航角相对偏差角，包括：
    获取所述飞行器的对地高度和所述飞行器的飞行高度；
    根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角。
  7. 根据权利要求6所述的方法，其特征在于，所述根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角，包括：
    根据所述飞行器的对地高度和所述飞行器的飞行高度，确定偏航角偏差相对偏移量；
    根据所述偏航角偏差角和所述偏航角偏差相对偏移量，确定所述偏航角相对偏差角。
  8. 根据权利要求7所述的方法，其特征在于，所述根据所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角偏差相对偏移量，包括：
    当满足以下条件中的任一项时，确定所述偏航角偏差角为所述偏航角偏差相对偏移量：
    所述飞行器的对地高度满足第一预设条件；
    所述飞行器的飞行高度满足第二预设条件时；以及
    所述偏航角偏差角的微分满足第三预设条件。
  9. 根据权利要求8所述的方法，其特征在于，所述第一预设条件为：
    所述飞行器的对地高度大于0.4m，且持续时间不小于0.5s。
  10. 根据权利要求8所述的方法，其特征在于，所述第二预设条件为：
    所述飞行器的飞行高度大于0.4m，且持续时间不小于0.5s。
  11. 根据权利要求8所述的方法，其特征在于，所述第三预设条件为：
    所述偏航角偏差角的微分的绝对值小于0.1且持续时间不小于0.5s。
  12. 根据权利要求8-11中任一项所述的方法，其特征在于，所述根据所述偏航角偏差角和所述偏航角偏差相对偏移量，确定所述偏航角相对偏差角，包括：
    当均满足以下条件时，确定所述偏航角误差相对补偿值为所述偏航角相对偏差角：
    所述飞行器的对地高度满足所述第一预设条件或当所述飞行器的飞行高度满足所述第二预设条件；
    所述偏航角偏差角的微分满足所述第三预设条件；且
    偏航角误差相对补偿值满足第四预设条件，其中，所述偏航角误差相对补偿值为所述偏航角偏差角与所述偏航角偏差相对偏移量的差。
  13. 根据权利要求12所述的方法，其特征在于，所述第四预设条件为：
    所述偏航角误差相对补偿值的绝对值小于0.1且持续时间不小于0.5s。
  14. 根据权利要求4所述的方法，其特征在于，所述根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量，包括：
    通过反馈控制算法对所述偏航角相对偏差角进行计算，以确定所述偏航角角速度补偿量。
  15. 一种飞行器的偏航角修正装置，应用于飞行器，其特征在于，所述装置包括：
    获取模块，用于获取IMU数据以及磁力计数据，其中，所述IMU数据包括IMU加速度信息和IMU角速度信息；
    确定模块，用于：
    根据所述磁力计数据，确定磁力计偏航角；
    根据所述磁力计偏航角，确定偏航角初始值；
    根据所述磁力计数据，确定偏航角角速度补偿量；
    根据所述IMU角速度信息以及所述偏航角角速度补偿量，确定修正后的角速度；以及
    根据所述修正后的角速度，确定偏航角相对值；
    融合偏航角生成模块，用于根据所述偏航角初始值以及所述偏航角相对值，生成融合偏航角。
  16. 根据权利要求15所述的装置，其特征在于，所述确定模块包括校准及坐标系转换模块，所述校准及坐标系转换模块用于：
    对所述磁力计数据进行校准，生成校准后的磁力计数据；
    获取所述飞行器的姿态角并根据所述飞行器的姿态角，生成旋转变换矩阵；
    利用所述旋转变换矩阵对所述校准后的磁力计数据进行坐标变换，以生成所述地面坐标系下的磁力计数据；
    根据所述地面坐标系下的磁力计数据，对比所述飞行器在当前位置下的标准磁场的磁力计数据，计算所述磁力计偏航角。
  17. 根据权利要求15所述的装置，其特征在于，所述确定模块还包括静止状态检测模块，所述静止状态检测模块用于：
    判断所述飞行器在当前时刻是否由静止状态变为运动状态；
    若是，则将所述磁力计偏航角作为所述偏航角初始值。
  18. 根据权利要求15所述的装置，其特征在于，所述确定模块还包括偏航角偏差判定及处理模块，所述偏航角偏差判定及处理模块用于：
    根据所述磁力计偏航角，确定偏航角偏差角；
    根据所述偏航角偏差角，确定偏航角相对偏差角；
    根据所述偏航角相对偏差角，确定所述偏航角角速度补偿量。
  19. 根据权利要求18所述的装置，其特征在于，所述偏航角偏差判定及处理模块用于：
    根据所述磁力计偏航角和上一时刻的融合偏航角，确定所述偏航角偏差角。
  20. 根据权利要求15所述的装置，其特征在于，所述偏航角偏差判定及处理模块用于：
    获取所述飞行器的对地高度和所述飞行器的飞行高度；
    根据所述偏航角偏差角、所述飞行器的对地高度和所述飞行器的飞行高度，确定所述偏航角相对偏差角。
  21. 根据权利要求20所述的装置，其特征在于，所述偏航角偏差判定及处理模块包括逻辑或运算模块，所述逻辑或运算模块用于：：
    当满足以下条件中的任一项时，确定所述偏航角偏差角为所述偏航角偏差相对偏移量：
    所述飞行器的对地高度满足第一预设条件；
    所述飞行器的飞行高度满足第二预设条件时；以及
    所述偏航角偏差角的微分满足第三预设条件。
  22. 根据权利要求21所述的装置，其特征在于，所述第一预设条件为：
    所述飞行器的对地高度大于0.4m，且持续时间不小于0.5s。
  23. 根据权利要求21所述的装置，其特征在于，所述第二预设条件为：
    所述飞行器的飞行高度大于0.4m，且持续时间不小于0.5s。
  24. 根据权利要求21所述的装置，其特征在于，所述第三预设条件为：
    所述偏航角偏差角的微分的绝对值小于0.1且持续时间不小于0.5s。
  25. 根据权利要求21-24中任一项所述的装置，其特征在于，所述偏航角偏差判定及处理模块包括逻辑与运算模块，所述逻辑与运算模块用于：
    当均满足以下条件时，确定所述偏航角误差相对补偿值为所述偏航角相对偏差角：
    所述飞行器的对地高度满足所述第一预设条件或当所述飞行器的飞行高度满足所述第二预设条件；
    所述偏航角偏差角的微分满足所述第三预设条件；且
    偏航角误差相对补偿值满足第四预设条件，其中，所述偏航角误差相对补偿值为所述偏航角偏差角与所述偏航角偏差相对偏移量的差。
  26. 根据权利要求25所述的装置，其特征在于，所述第四预设条件为：
    所述偏航角误差相对补偿值的绝对值小于0.1且持续时间不小于0.5s。
  27. 根据权利要求18所述的装置，其特征在于，所述确定模块包括反馈控制模块，所述反馈控制模块用于：
    通过反馈控制算法对所述偏航角相对偏差角进行计算，以确定所述偏航角角速度补偿量。
  28. 一种飞行器，其特征在于，包括：
    机身；
    机臂，与所述机身相连；
    动力装置，设于所述机身和/或所述机臂，用于为所述飞行器提供飞行的动力；以及
    飞行控制器，设于所述机身；
    其中，所述飞行控制器包括：
    至少一个处理器；以及，
    与所述至少一个处理器通信连接的存储器；其中，
    所述存储器存储有可被所述至少一个处理器执行的指令，所述指令被所述至少一个处理器执行，以使所述至少一个处理器能够执行权利要求1-14中任一项所述的方法。
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